JENGINEERIHG  LIBRARJ 


THE  AIRPLANE  ENGINE 


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THE 

AIRPLANE  ENGINE 


BY 

LIONEL  S.  MARKS,  B.Sc.,  M.M.E. 

PROFESSOR    OF   MECHANICAL   ENGINEERING,    HARVARD   UNIVERSITY 

MEMBER      AMERICAN      SOCIETY    MECHANICAL     ENGINEERS, 

FELLOW   AMERICAN   ACADEMY   ARTS   AND   SCIENCES 


FIRST  EDITION 


McGRAW-HILL  BOOK  COMPANY,  INC. 
NEW  YORK:  370  SEVENTH  AVENUE 

LONDON :  6  &  8  BOUVERIE  ST.,  E.  C.  4 

1922 


T '  l~7 
M  3 


Engineering 
Library 


COPYRIGHT,  1922,  BY  THE 
MCGRAW-HILL  BOOK  COMPANY,  INC. 


THE    MA.PUE    I'K  K  S  S    YORK 


MY     WIFE 

JOSEPHINE  PRESTON  PEABODY 


550891 


3 


Engineering 
Library 


COPYRIGHT,  1922,  BY  THE 
MCGRAW-HILL  BOOK  COMPANY,  INC. 


THE    MA.PLE    PRESS    YORK 


MY     WIFE 

JOSEPHINE  PRESTON  PEABODY 


550891 


PREFACE 

This  volume  attempts  two  things:  to  formulate  existing 
knowledge  of  the  functioning  of  the  airplane  engine  and  its 
auxiliaries;  and  to  present  and  discuss  the  essential  constructive 
details  of  those  engines  whose  excellence  has  resulted  in  their 
survival. 

The  material  here  collected  is  largely  new;  very  little  of  it 
could  have  been  written  before  the  war  and  only  a  small  frac- 
tion was  available  for  publication  before  1919.  It  is  based 
mainly  on  the  researches  and  engine  developments  originating 
during  the  war  and  resulting  from  the  war's  urgencies.  The 
researches  have  been  carried  out  almost  exclusively  under 
governmental  auspices;  in  the  United  States  at  the  Bureau  of 
Standards  and  at  the  Air  Service  experimental  plant  at  McCook 
Field;  in  Great  Britain  at  the  Royal  Aircraft  Factory  and  the 
National  Physical  Laboratory;  in  France  and  Germany  at 
equivalent  institutions.  Many  of  the  results  of  these  investiga- 
tions were  published  confidentially  during  the  war  in  Reports  of 
the  Bureau  of  Standards;  in  Bulletins  and  Technical  Orders  of 
the  Airplane  Engineering  Division  of  the  U.  S.  Army;  in  Reports 
of  the  (British)  Advisory  Committee  for  Aeronautics;  in  Bulle- 
tins de  la  Section  Technique  de  1'Aeronautique  Militaire;  and 
in  Technische  Berichte.  This  material  has  now  become  avail- 
able and  much  of  it  has  been  published  in  the  Reports  of  the 
(U.  S.)  National  Advisory  Committee  for  Aeronautics  and  in 
the  technical  press. 

Similarly,  the  constructive  details  of  most  of  the  existing 
airplane  engines  are  now  available,  chiefly  from  descriptions  of 
captured  machines.  The  German  and  Austrian  engines  captured 
by  the  British  were  subjected  to  a  technical  analysis  which  has  set 
a  new  standard  in  such  matters.  Not  only  were  the  engines  and 
their  auxiliaries  tested  exhaustively  for  performance  but  all  the 
parts  were  minutely  measured,  loads  and  stresses  calculated, 
and  the  metal  analyzed  for  chemical  composition.  The  French 
carried  out  similar  analyses  of  German  engines.  The  Germans 
published  corresponding,  though  less  detailed,  analyses  of 

vii 


viii  PREFACE 

French,  English  and  American  engines.  Since  the  war,  the 
U.  S.  Air  Service  has  also  analyzed  American  and  foreign  engines 
and  has  published  its  findings  in  Technical  Orders  and  Informa- 
tion Circulars. 

With  all  this  material,  the  designer  of  the  airplane  engine 
has  at  hand  more  detailed  precedent  from  which  to  depart  than 
is  available  for  other  types  of  engine. 

The  writer  desires  to  acknowledge  his  indebtedness  to  Professor 
E.  B.  Warner  and  to  Lieut.  E.  E.  Aldrin  for  assistance  in  obtaining 
information;  and  to  Mr.  R.  H.  Taylor  for  assistance  in  reading 
proofs  and  in  preparing  the  index. 

L.  S.  M. 

CAMBRIDGE,  MASS. 
January,  1922. 


CONTENTS 


PAGE 

PREFACE vii 

CHAPTER 

I.  POWER  REQUIRED  AND  POWER  AVAILABLE 1 

II.  ENGINE  EFFICIENCIES  AND  CAPACITIES  .    .    / 11 

III.  ENGINE  DYNAMICS 40 

IV.  ENGINE  DIMENSIONS  AND  ARRANGEMENTS 60 

V.  MATERIALS 114 

VI.  ENGINE  DETAILS 122 

VII.  VALVES  AND  VALVE  GEARS 151 

VIII.  RADIAL  AND  ROTARY  ENGINES 176 

IX.  FUELS  AND  EXPLOSIVE  MIXTURES 212 

X.  THE  CARBURETOR 245 

XI.  FUEL  SYSTEMS  . 289 

XII.  IGNITION.   .......  ^ .  295 

XIII.  LUBRICATION. 327 

XIV.  THE  COOLING  SYSTEM.   .    .    , 344 

XV.  GEARED  PROPELLER  DRIVES 378 

XVI.  SUPERCHARGING    . 386 

XVII.  MANIFOLDS  AND  MUFFLERS 416 

XVIII.  STARTING    .....;. .    .  424 

XIX.  POTENTIAL  DEVELOPMENTS 434 

INDEX  .  445 


IX 


THE  AIRPLANE  ENGINE 


CHAPTER  I 
POWER  REQUIRED  AND  POWER  AVAILABLE 

Power  Required  for  Flight. — An  airplane  in  flight  is  sustained 
by  the  lift  of  the  wings.  Consider  the  wing  as  a  thin  flat  plate, 
Fig.  1.  Four  forces  are  acting: 


FIG.  1. — Forces  acting  on  a  flat  plate. 

1.  The  weight,   W,   of  wing  and  parts  supported  thereby, 
downward. 

2.  The  thrust,   T,  or  forward  impulse  due  to  the  propeller. 

3.  The  lift,  L,  of  the  air  which  acts  in  a  direction  perpendicular 

1 


2  THE  AIRPLANE  ENGINE 

to  that  of  the  plane  with  respect  to  the  air,  and  produces 
sustentation. 

4.  The  wing  resistance  or  drag,  D,  measured  perpendicularly 
to  L,  the  component  of  the  total  force  on  the  wing  which  opposes 
forward  motion. 

At  constant  horizontal  speed,  L  =  W  and  D  =  T.  As  these 
pairs  of  forces  are  not  acting  in  the  same  lines  they  give  rise  to 
turning  moments  and  it  is  necessary  for  stability  that  L  X  m  = 
T  X  n.  Both  L  and  D  are  created  by  the  velocity  of  the  plane; 
the  direction  of  the  latter  is  concurrent  with  (but  opposite  to) 
the  path  of  flight.  The  former  is  directed  at  right  angles  with 
the  path  of  flight.  In  horizontal  flight  L  and  D  are  vertical  and 
horizontal  forces  respectively. 

Wing  Characteristics. — Flat  Plates. — A  plate  moving  with 
respect  to  the  air  undergoes  an  approximately  normal  pressure 
F  which  is  proportional  to  the  density  of  the  air  and  (within 
limits)  to  the  square  of  the  relative  velocity.  This  pressure  may 
be  resolved  into  components  L  and  D  perpendicular  and  parallel 
to  the  path  of  flight.  F2  =  L2  +  D2.  The  component  L  is 
useful,  while  D  is  objectionable.  The  pressure  F  depends  on  the 
area  of  the  plate,  but  varies  somewhat  with  its  shape. 

The  incidence,  i,  or  angle  between  the  plate  and  the  flight 
path  has  a  dominating  influence  on  the  resulting  pressure;  and 
particularly  on  the  relation  between  L  and  D.  Since  L  =  F  cos  i 
and  D  =  F  sin  i,  L/F  decreases  and  D/F  increases,  as  i  is 
increased  from  0  deg. 

For  a  given  angle  of  incidence  the  following  relations  hold: 

L  =  KLdAV2 
D  =  KDdAV2 

where  KL  and  KD  are  experimentally  determined  lift  and  drag 
coefficients  respectively;  d  is  the  air  density;  A  is  the  wing  area; 
and  V  the  relative  air  velocity.  The  usual  units  are  L  and  D  in 
pounds;  d,  relative  air  density  in  terms  of  normal  density;  A  in 
square  feet;  and  V  in  miles  per  hour.  From  the  above  equations 

T  7^- 

it  is  seen  that  T:=  ^-     That  is,  lift  is  obtained  with  minimum 

Lf          J\-D 

TV- 

drag  when^  is  a  maximum. 

&D 

It  is  found  that  more  favorable  ratios  of  KL/KD  are  obtained 
with  curved  wings,  as  in  Fig.  2,  than  with  flat  plates.  The  angle 


POWER  REQUIRED  AND  POWER  AVAILABLE  3 

of  incidence  of  such  a  wing  is  arbitrarily  defined  as  the  acute 
angle  between  the  wind  direction  and  the  lower  chord  of  the  wing. 
Figure  3  gives  the  values  of  KL)  KD,  and  KL/KD  or  L/D  for  the 
wing  or  aerofoil  section  shown  in  Fig.  2,  and  shows  a  maximum 
value  of  L/D  of  17  at  an  angle  of  incidence  of  3  deg. 


Chord-- 
FIG.  2.  —  Cross  section  of  wing. 


An  airplane  consists  not  only  of  the  wings  which  give  susten- 
tation,  but  also  of  other  members  such  as  the  fuselage,  radiator, 
landing  gear,  and  wing  bracing.  These  give  no  aid  in  sustaining 
the  plane  but  offer  a  resistance,  the  parasite  resistance,  which 


0.0028 
0.0024 
0.0020     20 
0.0016      16 
:*  0.00  12      12 
0.0008       8 
0.0004       4 
0       0 
-0.0004 
-nonnR 

•^-— 

•^\ 

0.0010 
0.0008 

0.0006 

a 
M 

0.0004 
0.0002 

o 

/ 

/ 

v 

/ 

/ 

f 

n 

x 

< 

/ 

1 

l 

x 

\i 

I/ 

x 

A 

^ 

% 

—  -*• 

*-" 

<K° 

^ 

/ 

/ 

Angle  of    Incidence  ,  Degrees 
FIG.  3. — Lift  and  drag  coefficients  of  a  wing. 

must  be  overcome  in  flight.     If  P  is  the  parasite  resistance  in 
pounds 

P  =  KNdV* 

where  KN  is  a  coefficient  with  practically  constant  value  in  any 
given  airplane. 


THE  AIRPLANE  ENGINE 

The  total  resistance,  R,  to  the  motion  of  the  plane  in  the  direc- 
tion of  flight  is  the  sum  of  the  wing  and  parasite  resistances,  or 

R  =  D  +  P  =  d(KDA  +  KN)V2 

This  resistance  must  equal  the  propeller  thrust,  T,  in  uniform 
flight.     The  horse  power  required  to  overcome  the  resistance  is 
H        1-467  flF    _  RV 
550         ~  375 

where  1.467  is  the  constant  to  convert  miles  per  hour  to  feet 
per  second  and  550  is  the  equivalent  of  1  hp.  in  foot-pounds  per 
second.  If  the  efficiency  of  the  propeller  is  e,  the  brake  horse 
power,  HB,  required  at  the  engine  for  horizontal  flight  at  speed  V 

RV     _d(KDA  +  KN)V* 
~  375e  375e 

The  lift,  L,  must  equal  the  weight  of  the  plane,  W,  in  horizontal 
flight,  and  the  equation  W  =  KLdAV2  must  be  satisfied  simul- 
taneously with  the  b.h.p.  equation,  with  values  of  KL  and  KD 
corresponding  to  some  one  angle  of  incidence.  In  a  design 
Wj  A,  V,  KN  are  known  by  assumption  or  calculation  and  it  is 
required  to  find  HR,  KL  and  KD  for  a  series  of  values  of  F.  Since 
KL  and  KD  are  functions  of  i  there  are  only  two  independent 
variables  HB  and  i  to  be  found. 

The  power  required  in  an  airplane  of  200  sq.  ft.  wing  area 
with  a  total  weight  of  1,200  Ib.  using  the  wing  with  the  properties 
shown  in  Fig.  3  and  with  parasite  resistance  P  =  0.0025  F2 
is  shown  in  Fig.  4.  It  will  be  seen  that  the  plane  has  a  minimum 
possible  velocity  of  about  44  miles  per  hour  but  that  the  power 
required  to  drive  it  starts  to  increase  very  rapidly  if  the  speed 
gets  below  46  miles  per  hour.  The  flying  conditions  are  for 
speeds  of  46  miles  per  hour  or  higher.  At  lower  speeds  there  are 
reversed  controls,  that  is,  more  power  is  required  to  go  slower, 
and  the  angle  of  incidence  is  so  high  that  the  plane  will  be  in 
danger  of  stalling.  The  power  required  curve  of  Fig.  4  is  of  a  form 
which  may  be  regarded  as  typical. 

If  the  propeller  does  more  work  than  is  required  the  plane  will 
climb  and  its  rate  of  climbing  will  depend  only  on  the  amount  of 
the  excess  power;  all  the  additional  work  goes  into  raising  the 
plane.  If,  on  the  other  hand,  power  is  deficient  and  the  propel- 
ler does  less  work  than  is  required  to  keep  the  plane  in  level 
flight  at  the  existing  speed,  the  plane  will  glide  down  and  the  work 


POWER  REQUIRED  AND  POWER  AVAILABLE 


done  by  the  falling  plane  (weight  X  fall)  will  be  exactly  equal 
to  the  difference  between  the  work  required  to  keep  the  plane 
moving  with  its  existing  speed  in  the  direction  of  flight  and  the 
work  done  by  the  propeller.  If  the  power  available  is  at  any 
time  in  excess  of  that  required  for  level  flight,  it  can  be  reduced  by 
throttling  the  engine.  There  is  only  one  speed  possible  in 
level  flight  for  any  given  angle  of  incidence.  Opening  the  throttle 
does  not,  as  with  automobiles,  increase  the  speed  in  level  flight, 
but  will  start  the  airplane  climbing  unless  the  incidence  is  also 


/ 

100 

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Velocity  ,   Miles   per    Hour 
FIG.  4.  —  Characteristics  of  an  airplane  in  level  flight. 

changed  by  operating  the  elevator.  The  power  supplied  by  the 
engine  with  a  given  incidence  cannot  affect  the  velocity  of  the  plane 
but  will  determine  whether  the  plane  climbs,  flies  level,  or  glides. 

TV 

The  horse  power  available  for  driving  the  plane  is  given  by 


that  is,  at  any  plane  speed  it  depends  only  on  the  propeller 
thrust.  As  the  thrust  depends  on  the  engine  speed  both  the 
engine  and  the  propeller  characteristics  must  be  examined  before 
the  thrust  can  be  determined. 

Air  Propellers.  —  The  propeller  converts  the  torque  at  the 
engine  crankshaft  into  a  thrust  along  the  axis  of  the  propeller 
shaft.  Its  action  is  remotely  similar,  to  the  turning  of  a  solid 


6  THE  AIRPLANE  ENGINE 

screw  in  a  solid  nut  and  the  same  general  terminology  is  em- 
ployed. The  axial  distance  that  would  be  travelled  by  the  pro- 
peller for  one  revolution  if  the  air  were  incompressible  is  called 
its  nominal  pitch;  the  nominal  pitch  divided  by  the  outside 
diameter  is  the  nominal  pitch  ratio.  An  air  propeller  does  not, 
however,  advance  a  constant  amount  for  one  revolution;  it  is 
advancing  in  a  medium  which  is  readily  displaced,  and  it  is 
found  that  it  must  advance  each  revolution  a  distance  greater 
than  the  nominal  pitch  if  it  is  not  to  displace  the  air  backward. 
This  greater  distance  is  called  the  dynamic  pitch.  The  difference 
between  the  dynamic  pitch  and  the  actual  or  effective  pitch  is  called 
the  slip.  Positive  thrust  can  be  obtained  only  when  there  is  slip 
and  the  greater  the  slip  the  greater  will  be  the  thrust;  maximum 
thrust  is  obtained  when  the  plane  is  on  the  ground  in  which  case 
the  effective  pitch  is  zero. 

If  a  propeller  makes  N  revolutions  per  minute  and  the  plane 
moves  through  the  air  with  a  velocity  of  V  feet  per  minute  the 
effective  pitch  is  V/N  and  the  effective  pitch  ratio  is  V/ND. 

The  characteristics  of  propellers  are  most  easily  determined  by 
tests  on  models.  It  is  found  that  geometrically  similar  propellers 
have  the  same  characteristics  for  the  same  values  of  V/ND. 
The  important  characteristics  are  the  thrust,  T,  lb.;  torque, 
Q,  lb.-ft.;  torque  horse  power,  HB;  thrust  horse  power,  HR;  and 
efficiency,  e.  The  torque  horse  power  is  necessarily  the  same  as 
the  engine  brake  horse  power.  The  efficiency,  e  =  HR/HB. 

Torque  and  thrust  are  given  by  the  following  equations: 

w  V2  D3  -  a 
Q  = 


T  = 


1000 
w  V*  D*  •  b 


100 

where  a  and  b  are  the  torque  and  thrust  coefficients  respectively 
and  w  is  the  air  density  in  pounds  per  cubic  foot. 

Efficiency  is  obtainable  directly  from  these  coefficients. 

Efficiency  = 

Thrust  horse  power  =    TV     =  IQfr    V_   =         b     _7_ 
Torque  horse  power     2irQN      2ira  ND  a    ND 

Values  of  these  coefficients  and  of  e  have  been  determined  by 
Durand1  for  propellers  of  many  types  and  proportions.     The 
1  Nat.  Adv.  Comm.  on  Aeronautics,  1917. 


POWER  REQUIRED  AND  POWER  AVAILABLE 


curves  of  Fig.  5  give  typical  values  obtained  from  three  propel- 
lers which  differ  only  in  nominal  pitch  ratio;  the  values  of  the 
pitch  ratio  are  0.5,  0.7  and  0.9  respectively.  By  the  use  of  such 
curves  the  horse  power  required  can  be  easily  obtained.  For 
example,  assume  a  propeller  8  ft.  in  diameter  (Fig.  5)  with  pitch 
ratio  0.9,  making  1,200  r.p.m.  and  with  a  speed  of  72  miles  per 
hour  (105.6  ft.  per  second)  at  an  elevation  of  3,000  ft.  The  value 


0.1      0.2     03     0.4     0.5     0.6     0.1     0.8     0.9     1.0 


i      12 


FIG.  5. — Propeller  coefficients  and  efficiencies. 


of  V/ND  =  0.66.     From  Fig.  5,6  =  0.685,  and  a  =  0.970.     Also 
w  =  0.071  Ib.  per  cubic  foot.     Then 

T  =  0.071  X  (105.6)  2X  82  X  0.685 

J.UU 


Q  = 


0.071  X  (105.6)2  X  83  X  0.970 


1,000 
Torque  horse  power,  HB,  = 

2irQN    =  2r  X  393  X  1,200 
33,000  " 


=  89.5 


33,000 
Efficiency  =  1.59  X  g^  X  0.66  =  0.742 


8  THE  AIRPLANE  ENGINE 

The  efficiency  can  also  be  read  directly  from  Fig.  5. 

Power  Available  for  Flight.  —  The  horse  power  available  with  a 
given  engine  and  given  plane  speed,  F,  can  now  be  determined. 
Throughout  a  considerable  range  of  revolutions  per  minute  the 
mean  effective  pressure  in  the  engine  is  practically  constant  but 
falls  off  at  highest  speeds.  If  it  is  assumed  constant  the  propeller 
torque  will  be  constant.  The  torque  equation  can  be  written 

/  V  y  =     Q 


1,000     \ND/ 
or  since  Q  and  D  are  constant 

Constant 


N  = 


F-x 


(ND) 


y 

Assuming  various  values  of   -^jj  and  finding  in   Fig.  5  the 

corresponding  values  of  a,  a  series  of  values  of  N  can  be  obtained 
and  substituting  these  values  in  V/ND  the  corresponding  values 
of  V  are  determined.  That  is,  the  change  of  revolutions  per 
minute  with  flying  speed,  F,  or  with  V/ND  is  obtained.  In 
Fig.  6  there  are  plotted  curves1  showing  this  variation  for 
propellers  of  nominal  pitch  ratios  of  0.5,  0.7  and  0.9  respectively. 
These  curves  do  not  correspond  exactly  to  the  coefficients  of  Fig. 
5;  they  average  the  results  of  Durand's  tests  and  apply  fairly 
well  to  standard  forms  of  propeller.  The  unit  values  of  V/ND 
and  N  are  those  corresponding  to  maximum  efficiency.  The 
product  of  the  ordinates  and  abscissae  for  any  point  is  (V/ND) 
X  N  =  V/D,  and  since  D  is  constant,  the  ratio  of  this  product 
at  two  points  is  also  the  ratio  of  the  corresponding  plane 
velocities. 

As  an  example  suppose  a  propeller  with  nominal  pitch  ratio 
0.7  designed  for  an  airplane  flying  normally  at  100  miles  per  hour 
and  that  the  full  engine  power  is  absorbed  at  1,600  r.p.m.  If 
the  speed  of  the  machine  increases  so  that  V/ND  =  1.1,  then 
N  becomes  1.052  and  V  increases  in  the  ratio  1.1  X  1.052  = 
1.157.  The  propeller  will  then  turn  1.052  X  1,600  =  1,683  r.p.m. 
on  full  throttle  at  a  plane  speed  of  115.7  miles  per  hour. 

The  torque  horse  power  is  proportional  to  the  engine  speed; 
the  efficiency  can  be  obtained  from  Fig.  5,  and  multiplied  by 

1  Supplied  by  E.  P.  Warner. 


POWER  REQUIRED  AND  POWER  AVAILABLE 


the  torque  horse  power  will  give  the  available  horse  power.  In 
the  example  just  given,  assume  that  the  propeller  has  the  char- 
acteristics shown  in  Fig.  5  and  that  the  torque  horse  power  at 
maximum  efficiency  is  100.  Maximum  efficiency  is  seen  to  be 
for  V/ND  =  0.64,  which  therefore  corresponds  to  the  unit  of 
Fig.  6.  For  V/ND  =  1.1  on  Fig.  6  we  have  V/ND  =  0.64  X 
1.1  =  0.704  in  Fig.  5;  the  corresponding  efficiency  is  seen  to  be 
0.74.  The  torque  horse  power  is  100  X  1.052  and  the  available 
horse  power  is  105.2  X  0.74  =  78.  By  finding  the  available 


1,3 


1.2 


0.9 


\/ 


0.4 


0.6 


0.8 


1.0 

V/ND 


1.2 


1.4 


FIG.  6. — Variation  of  revolutions  per  minute  N,  with  speed  of  airplane,   V; 
engine  torque  constant. 

horse  power  for  a  number  of  other  values  of  V/ND  and  plotting 
them  against  V  the  available  horse  power  curve  of  Fig.  4  is 
obtained. 

If  engine  torque  varies  with  the  engine  speed,  N,  another  pro- 
cedure must  be  followed  in  finding  the  values  of  N  corresponding 
to  different  values  of  V.  Taking  the  propeller  torque  equation 

Q  =  ^j-nrvr  a  series  of  curves  may  be  drawn,  as  in  Fig.  7,  each 

IjUUU 

giving  the  change  in  propeller  torque  with  speed  at  constant  V.  If 
the  engine  torque  is  now  drawn  on  the  same  figure  its  intersection 


10 


THE  AIRPLANE  ENGINE 


with  the  constant  V  curves  will  give  the  values  of  N  at  which 
engine  and  propeller  torques  are  equal,  that  is,  it  will  give  the 
operating  speeds.  The  available  horse  power  can  then  be 
found  by  the  use  of  the  efficiency  curve  of  Fig.  5 

Returning  to  Fig.  4  it  is  seen  that  with  wide-open  throttle  the 
available  horse  power  is  equal  to  the  required  horse  power  at  45 
miles  per  hour  and  at  90  miles  per  hour;  between  these  limits  the 
available  horse  power  is  in  excess  of  the  power  required  for  level 


TOO 


100. 


1000 


1100 


1200 


1300 


1400          1500 
R.  P.  M 


1600 


noo 


1800         190C 


FIG.  7. — Variation  of  engine  and  propeller  torque  with  revolutions  per  minute 
at  various  airplane  speeds. 

flight;  outside  these  limits  the  engine  power  is  insufficient  and  the 
plane  will  glide.  Between  45  and  90  miles  per  hour  level  flight 
can  be  maintained  only  by  closing  the  throttle  and  the  speed  of 
flight  will  depend  on  the  angle  of  incidence.  With  the  throttle 
wide  open  the  plane  will  climb  and  its  rate  of  climb  will  be  great- 
est at  that  speed  and  corresponding  angle  of  incidence  at  which 
the  difference  between  the  available  power  and  required  power 
is  a  maximum.  In  the  case  of  Fig.  4  this  will  be  at  about  65 
miles  per  hour  and  an  angle  of  incidence  of  3.6  deg. 


CHAPTER  II 


ENGINE  EFFICIENCIES  AND  CAPACITIES 

A  typical  airplane  engine  is  shown  in  transverse  section  in 
Fig.  8  and  in  longitudinal  section  in  Fig.  9.     It  has  six  vertical, 
single-acting,  water-cooled  cylinders  in  a  row,  driving  the  crank- 
shaft, b,  through  pistons,  p, 
piston    pins,    g,    connecting 
rods,    r,   and   crankpins,    k. 
The  crankshaft  is  supported 
on    seven    bearings,    carries 
the  propeller  hub,  d,  at  its 
forward   end  with  a  thrust 
bearing,  e,  behind  it  and  has 
a  bevel  wheel,  /,  at  its  rear 
end   meshing    with   another 
bevel  wheel  on  the  vertical 
shaft,  Z.     The  shaft  I  drives 
the    camshaft,    c,    through 
bevel   gearing   at  its  upper 
end  and  carries  at  its  lower 
end    the    centrifugal    water 
pump,  7i,  and  the  two  gear 
oil  pumps,  o;  it  also  drives 
the    magneto,    ra,    through 
helical  gears,  q.     The  cam- 
shaft, c,  carries  inlet   cams 
which    act    on    the    rocker 
levers,  t,  and  open  the  inlet 
valves,   v,  against  the   com- 
pression of  the  valve  springs, 
h;  there  is  one  inlet  valve  to 
each  cylinder.     The  cam- 
shaft also  carries  cams  which 
act  directly  on  the  tappets  of  the  exhaust  valves,  u,  of  which 
there  are  two  per  cylinder.     The  water  jackets,  w,  surround  the 
cylinders  and  the  valve  cages.     Air  enters  the  carburetor,  a,  and 
goes  through  the  inlet  manifold,  i,  and  past  the  inlet  valve,  v,  to 

11 


FIG. 


8. — Transverse   section   of   Siddeley 
"Puma"  airplane  engine. 


12 


THE  AIRPLANE  ENGINE 


ENGINE  EFFICIENCIES  AND  CAPACITIES 


13 


Pressures,Lb  perSq.. 

—  ro  OJ  -1 

0  0  0  C 
O  0  0  0  C 

d 

\ 

\ 

\ 

*^ 

c 

V 

x 

^ 

••» 

^= 

q 

Volumes 

FIG.  10. — Ideal  indicator  card 
of  Otto-cycle  engine. 


the  cylinder,  x.  After  compression  the  charge  is  ignited  by  the 
spark  plugs,  s,  which  get  their  current  from  the  magneto,  m. 

Actions  Occurring  in  the  Cylinder. — Figure  10  is  a  theoretica 
indicator  card  for  an  engine  using  the  Otto  cycle,  which  is  always 
used  in  aviation  engines.  The  indicator  card  shows  (vertically) 
the  gas  pressure  inside  the  cylinder  at  each  position  of  the  piston. 
At  the  position  a  the  piston  is  at  the  end  of  its  stroke  most  remote 
from  the  crankshaft  and  the  volume  of  is  the  volume  of  the  clear- 
ance or  combustion  space;  the  total  volume  displaced  by  the 
piston  is  represented  by  fg  and  is  the 
product  of  the  piston  area  by  its  stroke; 
the  maximum  volume  of  gas  in  the  cyl- 
inder is  og. 

The  cycle  of  operations  inside  the 
cylinder  begins  with  the  piston  at  a 
and  with  the  gas  inlet  valve  open, 
establishing  free  communication  be- 
tween the  cylinder  and  the  external 
air.  The  piston  makes  its  stroke  from 
a  to  b  with  the  inlet  valve  kept  open; 

the  pressure  will  remain  atmospheric.  This  is  the  suction  stroke. 
The  inlet  valve  now  closes  and  the  piston  returns.  Since  both 
valves  are  closed,  the  mixture  in  the  cylinder  is  compressed  (curve 
be,  compression  stroke)  the  final  pressure  depending  mainly  upon 
the  volume  of  the  clearance  space  into  which  the  gas  is  crowded  at 
the  inner  end  of  the  stroke. 

When  the  piston  is  at  c  (actually  a  little  sooner  than  this) 
the  compressed  mixture  is  ignited  and  the  pressure  suddenly 
increases  (line  cd).  This  pressure  drives  the  piston  back,  the 
burnt  mixture  expanding  (curve  de)  while  the  valves  remain 
closed.  This  is  the  power  or  expansion  stroke. 

At  the  end,  e,  of  the  expansion  stroke,  the  exhaust  valve  opens. 
The  pressure  drops  to  atmospheric  (line  eb).  The  piston  returns 
and  the  burned  gas  is  driven  out  at  the  exhaust  valve  at  atmos- 
pheric pressure .  At  the  end  of  this  exhaust  stroke ,  ba,  the  exhaust 
valve  closes,  the  inlet  valve  opens,  and  the  cycle  recommences. 

SUMMARY 

STROKE  ACTION  INLET  VALVE       EXHAUST  VALVE 

First  out  Suction  Open  Closed 

First  in  Compression  Closed  Closed 

Second  out  Explosion  and  expansion  Closed  Closed 

Second  in  Exhaust  Closed  Open 


14  THE  AIRPLANE  ENGINE 

Pressures  and  Temperatures  in  the  Ideal  Cycle. — Let  p  be 

the  absolute  gas  pressure  in  pounds  per  square  inch,  v  the  volume 
in  cubic  feet,  t  the  temperature  Fahrenheit  and  T  the  absolute 
temperature.  The  suffixes  a,  6,  c,  etc.,  indicate  the  points  on  the 
cycle,  shown  in  Fig.  10,  at  which  these  quantities  are  being 
considered.  In  the  ideal  cycle,  the  volume  vb  —  va  of  explosive 
mixture  is  taken  into  the  cylinder  during  the  admission  period 
and  is  still  at  atmospheric  pressure  and  temperature  at  the  end 
of  the  stroke  6.  The  mixture  or  " charge"  is  now  compressed 
adiabatically,  that  is,  without  addition  or  abstraction  of  heat. 
As  a  result  of  the  compression  the  pressure  and  temperature 
increase;  the  compression  pressure  is  given  by  the  equation 

pb       \vj 

where  r  =  —  is  the  ratio  of  compression.  The  temperature 
is  given  by 


T 

_c    —    r0.4 

Tb~ 

If  the  atmospheric  pressure,  pb,  is  14.7  Ib.  per  square  inch  and 
the  temperature  tb  is  60°F.,  or,  Tb  =  460  +  60  =  520°  abso- 
lute, then,  with  r  =  4.8,  pc  =  9  X  14.7  =  132.3  Ib.  per  square 
inch  and  Te  =  1.873  X  520  =  973°  absolute,  or,  tc  =  973  - 
460  =  513°F. 

As  a  result  of  the  explosion  at  c  the  heat  of  combustion  is  lib- 
erated and  is  used  in  heating  up  the  charge.  The  heat  of  com- 
bustion may  be  assumed  to  be  80  B.t.u.  per  cubic  foot  of 
charge  admitted,  or  80  (vb  —  va)  B.t.u.  The  weight,  w,  of  gas 
in  the  cylinder  is  given  by  the  perfect-gas  equation 

144  pv  =  WRT, 

where  R  is  the  gas  constant,  which  may  be  assumed  to  have  the 
value  52  for  the  usual  explosive  mixture.  For  the  conditions 
assumed,  the  weight  of  gas  is 

144  pv      144  X  14.7 


52  X  520 


=  °'°784  *' 


The  heat  required  to  raise  1  Ib.  of  this  gas  1°F.  while  the  volume 
remains  unchanged  is  called  the  specific  heat  at  constant  volume, 
CVJ  and  may  be  taken  as  0.171  B.t.u.  The  rise  in  temperature 
during  explosion  is  given  by  the  equation 


ENGINE  EFFICIENCIES  AND  CAPACITIES 


15 


Heat  of  explosion,  H,  =  weight  of  gas  X  specific  heat  X  rise  of 
temperature,  or 

H  =  wCv(Td  -  Tc) 

For  the  conditions  assumed, 

80(z^-  Va)  =  0.0784  vbCv(Td  -  Tc} 
and  since 


=  4.8 

Va 


~  T<  -  80 


.171   -  4'74° 


and  the  explosion  temper  aturel_T^l  =  4,740  +  973  =  5,713°. 
The  explosion  pressure,  pd,  is  given  by  the  equation 


or,  pd  = 


5,713 
973 


P 


Tc 


X  132.3  =  778  Ib.  per  square  inch. 


The  expansion  curve  in  the  ideal  cycle  is  adiabatic,  so  'that  the 
equations  are  the  same  as  for  the  compression.  The  pressure 
pe  at  release  or  beginning  of  exhaust,  e,  is  given  by 


=   rl-4 


and 


and 


Pe 

=    °-« 


=  r 


778 


Consequently,  pe  =  -TT-  =  86.4  Ib.  per  square  inch 
y 


Te  = 


5,713 


3,060C 


:400 


300 


200 


100 


1.873 

The  actual  indicator  card  is  shown 
in  Fig.  11.  At  the  end  of  the  exhaust 
stroke  the  pressure  in  the  cylinder  is 
above  that  of  the  atmosphere;  usually, 
pa  =  15.4  to  17  Ib.  per  square  inch. 
This  pressure  falls  as  the  piston 
moves  down  and  the  fresh  charge  is 
drawn  in.  Mixing  with  the  residual 
burned  gas  this  charge  is  somewhat  FlG-  ,11;~^ctua}  ind!cator 

.  card  of  Otto-cycle  engine. 

heated,  the  temperature,  fo,  being  usu- 
ally from  170  to  260°  and  the  pressure,  pb,  12  to  14  Ib.  absolute. 
This  pressure  depends  on  the  engine  speed  and  on  the  resistance 


Exhaus 
q. 


Admission- 

Volumes 


16 


THE  AIRPLANE  ENGINE 


encountered  by  the  fresh  charge  in  the  carburetor,  manifold, 
and  inlet  valve. 

The  compression  curve  be  is  not  adiabatic,  but  may  be 
represented  by  the  expression  pcvcn  =  pbVbn,  where  n  has  a  value 
between  1.23  and  1.35.  The  compression  pressure,  pc,  is  deter- 
mined by  the  clearance  volume:  110  to  120  Ib.  gage  is  near  the 
average  of  good  practice  for  water-cooled  engines.  The  com- 
pression pressure  tends  to  fall  with  increase  of  engine  speed  in 

consequence  of  the  increase 
of  vacuum  in  the  cylinder 
at  the  moment  of  closure  of 
the  inlet  valve.  The  curve 
for  the  Liberty  engine  (Fig. 
12)  shows  this  effect.  The 
maximum  temperature,  t*, 
following  ignition,  is  usu- 
ally from  2,500  to  3,200°F.,  and  the  maximum  pressure,  pd  be- 
tween 300  and  400  Ib.  When  pc  is  increased,  pd  also  increases. 
The  expansion  curve  de  also  is  not  adiabatic  but  may  be  rep- 
resented by  pdvdn  =  peven,  in  which  n  =  1.27  to  1.5;  the  low  values 
are  realized  in  cases  where  combustion  continues  after  expansion 
has  been  well  started.  The  terminal  pressure,  pe,  is  from  38  to 
75  Ib.,  and  te  from  1,200  to  2,000°F.  The  pressure  during  exhaust 
varies  from  17  to  15.4  Ib.  per  square  inch. 

The  compression  ratio,  —>  is  fixed  by  the  clearance  and  may  be 

Vc 

varied  in  a  given  engine  by  the  use  of  pistons  with  heads  of  dif- 


1200         1400        1600         1800        2000 
Engine  Revolutions   per  Minu-fce 

FIG.  12. — Variation  of  compression  pressure 
with  engine  speed. 


TABLE  OF  COMPRESSION  PRESSURES,  POUNDS  PER  SQUARE  INCH 


Compression  ratios 


n 

3.5 

4.0 

4.5 

5.0 

6.0 

7.0 

1.25 

59.9 

70.6 

81.9 

93.2 

116.5 

143.0 

1.30 

63.7 

75.8 

88.3 

101.3 

128.4 

156.9 

1.35 

67.8 

81.2 

95.2 

109.8 

140.4 

172.9 

1.41 

73.1 

88.3 

104.2 

120.9 

156.4 

194.3 

ferent  shapes.     It  is  the  chief  factor  in  determining  pc.    The 
accompanying  table  is  based  on  pb  =  12.5.     A  high  compression 


ENGINE  EFFICIENCIES  AND  CAPACITIES  17 

ratio  increases  power  output  and  efficiency.     It  also  increases 
the  temperature  at  the  end  of  compression  since 


=  (!l\  n 

Tb      tb  +  460    "  \pj 

A  high  value  of  tc  may  cause  preignition  and  thus  fixes  a  limit  of 
compression  ratio  which  must  not  be  exceeded.  Usual  values  are 
from  4.5  in  hydroplanes  to  5.6  for  high-altitude  land  machines. 
Values  of  compression  ratio  are  given  for  various  engines  on  pages 
66  to  70,  where  it  is  seen  that  high  values  result  in  increased 
power  output  per  unit  of  cylinder  volume.  Values  of  six  or  higher 
may  be  used  for  high  altitude  work,  but  the  engines  will  develop 
preignition  if  operated  at  full  throttle  near  the  ground.  When 
operated  on  partial  throttle  the  entering  charge  is  of  reduced 
weight,  both  because  of  lower  pressure  and  because  of  dilution 
of  the  fresh  mixture  with  a  relatively  greater  weight  of  burnt  gas 
remaining  over  from  the  previous  cycle.  Consequently  the  heat 
developed  per  cycle  is  less  and  the  mean  temperatures  in  the 
cylinders  are  reduced. 

Efficiency. — If  the  working  substance  in  the  cylinder  followed 
the  laws  of  a  perfect  gas 

pv  =  RT 
and  Cv  =  constant 

and  if  the  combustion  were  instantaneous  and  complete,  the 
efficiency  of  the  cycle  would  be  equal  to 

-i 

(i) 

where  r  is  the  compression  ratio.  It  is  here  assumed  further 
that  the  cycle  takes  place  without  any  heat  exchange  between 
the  working  charge  and  the  cylinder.  The  efficiency  so  found  is 
the  highest  possible  efficiency  for  an  engine  operating  on  the  Otto 
cycle  and  could  be  attained  only  under  the  conditions  stated 
above.  It  is  sometimes  called  the  air -cycle  efficiency.  Its  value 
for  various  compression  ratios  is  given  in  the  following  table : 

r=3.50     4.00    4.50     5.00     5.50     6.00     6.50     7.00 
e  =  0.40     0.43     0.45     0.48    0.50    0.51     0.53     0.54 

The  efficiencies  actually  obtained  in  airplane  engines  are  seldom 
greater  than  60  per  cent  of  these  values.  For  instance,  with 
r  =  5.5,  the  actual  efficiency  will  not  exceed  0.6  X  0.5  =  0.3 


18  THE  AIRPLANE  ENGINE 

and  will  in  general  be  between  0.25  and  0.3.  This  considerable 
discrepancy  between  the  actual  performance  of  the  engine  and 
the  air  cycle  efficiency  is  due  to  a  variety  of  causes,  the  principal 
of  which  are  as  follows : 

I.  The  theoretical  cycle  assumes  that  the  total  heat  (lower 
heat  value)  of  combustion  of  the  fuel  taken  into  the  cylinder  is 
utilized  during  explosion  in  heating  up  the  working  mixture. 
This  is  not  actually  the  case  for  two  reasons : 

(a)  The  whole  of  the  heat  of  combustion  is  not  evolved  during 
explosion  because  combustion  is  not  instantaneous,  so  that 
combustion  will  continue  for  part  (or  with  incorrect  mixture, 
for  the  whole)  of  the  expansion  stroke,  thereby  reducing  the 
amount  of  heat  available  for  conversion  into  work.  Furthermore, 
complete  combustion  at  the  end  of  explosion  is  not  attainable 
because  chemical  equilibrium  requires  the  presence  of  a  certain 
amount  of  hydrogen  and  carbon  monoxide.  Their  existence  is 
commonly  ascribed  to  dissociation.  The  amount  of  heat  suppres- 
sion from  these  causes  is  not  considerable  in  a  high-grade  engine 
operating  with  gasoline  and  with  a  good  mixture. 

(6)  Some  of  the  heat  actually  evolved  goes  to  the  cylinder 
walls  by  radiation  and  conduction.  The  total  heat  so  going 
to  the  walls  in  airplane  engines  is  from  25  to  30  per  cent  of  the 
total  heat  of  combustion.  If  the  heat  were  abstracted  from  the 
burning  mixture  during  the  explosion  it  would  result  in  a  loss 
of  efficiency  of  the  same  magnitude.  The  actual  passage  of 
heat  from  the  mixture  to  the  walls  continues  from  the  middle 
of  compression  to  the  end  of  exhaust.  Throughout  the  whole  of 
this  time  the  gases  are  hotter  than  the  walls.  The  heat  flows 
in  the  opposite  direction  during  the  admission  period  and  the 
first  part  of  the  compression,  but  the  amount  of  heat  thus  flowing 
is  small,  as  the  temperature  difference  between  the  walls  and  the 
gases  is  small.  Such  heat  as  passes  off  during  the  explosion 
and  the  first  part  of  the  expansion  stroke  may  be  regarded  as 
entirely  lost  to  the  engine;  the  heat  flow  to  the  walls  near  the 
end  of  expansion  and  during  exhaust  is  no  loss  at  all,  as  it  is 
necessary  to  discharge  the  hot  gases  and  it  is  immaterial,  from 
the  point  of  view  of  efficiency,  whether  the  heat  is  carried  away 
by  the  jacket  water  or  in  the  exhaust  gases. 

General  experience  would  indicate  that  more  than  one-half 
of  the  total  heat  given  to  the  jacket  may  be  regarded  as  abstracted 
by  radiation  or  conduction  from  the  working  substance  during 


ENGINE  EFFICIENCIES  AND  CAPACITIES  19 

explosion  and  the  early  part  of  expansion.  It  should  be  noted 
in  this  connection  that  the  heat  of  the  jacket  water  includes 
most  of  the  friction  work  between  the  piston  and  cylinder,  which 
is  a  considerable  fraction  of  the  total  friction  of  the  engine.  As 
the  jacket  heat  is  25  to  30  per  cent,  the  heat  lost  during  explosion 
by  radiation  and  conduction  may  be  taken  as  not  more  than  12 
to  15  per  cent  of  the  heat  of  complete  combustion. 

II.  The  working  substance  is  not  a  perfect  gas  and,  in  particu- 
lar, it  is  not  true  that  the  specific  heat  at  constant  volume  is  a 
constant.  It  is  found  on  investigation  that  the  gases  (C02,  N£, 
H20  etc.)  which  are  present  in  the  cylinder  after  explosion  have 
specific  heats  which  increase  considerably  with  increase  of  tem- 
perature. These  specific  heats  follow  the  equation 

Cv  =  a  +  bt 

where  a  and  6  are  constants. 

The  efficiency  of  the  cycle  is  diminished  as  a  result  of  this 
increase  of  specific  heat.  The  immediate  result  is  that  the  rise 
of  temperature  during  explosion,  for  a  given  amount  of  fuel 
burned,  is  diminished  and  consequently  the  pressure,  pd  (Fig. 
10),  is  lower  than  would  be  realized  with  constant  specific  heat. 
The  expansion  curve  de  is  consequently  lowered  and  the  work 
of  the  cycle  diminished.  The  efficiency  with  adiabatic  expan- 
sion and  compression  but  with  variable  specific  heat  is  given 
by  the  expression,1 


where  e  is  the  air-cycle  efficiency  and  Td  and  Tj>  are  the  abso- 
lute temperatures  at  the  end  of  explosion  and  beginning  of 
compression  respectively.  The  constants  a  and  b  have  values 
of  about  0.194  and  0.051  X  10~3  for  the  average  working  mix- 
ture. For  the  conditions  customarily  met  in  airplane  engines 

E 

the  ratio  of  the  two  efficiencies  —  is  about  0.80;  in  other  words, 

6 

the  theoretical  efficiency  with  the  actual  working  substance  is 
only  80  per  cent  of  that  which  would  be  attainable  if  these  sub- 
stances were  perfect  gases. 

The  actual  working  substance  consists  almost  exclusively  of 
nitrogen,  water  vapor  and  carbon  dioxide.  All  three  of  these 
substances  show  a  considerable  increase  in  specific  heat  with  rise 

1  WIMPERIS,  The  Internal  Combustion  Engine,  p.  85. 


20 


THE  AIRPLANE  ENGINE 


in  temperature  and  the  last  two  dissociate  at  high  temperatures, 
especially  at  low  pressures.  The  following  tables1  give  mean 
specific  heats  at  constant  volume,  and  percentage  dissociation. 


MEAN  SPECIFIC  HEATS  AT  CONSTANT  VOLUME   (IN  B.T.U.  PER   DEGREE 
FAHRENHEIT)  BETWEEN  200°F.  AND  THE  STATED    TEMPERATURE 


Temperature,  degrees 
Fahrenheit 

930 

1;830 

2,730 

3,630 

4,530 

5,430 

Nitrogen.   . 

0  185 

0  188 

0  196 

0  205 

0  214 

0  225 

Water  vapor  

0.350 

0.385 

0.425 

0.468 

0.540 

0.623 

Carbon  dioxide  

0.187 

0.217 

0.229 

0.238 

0.247 

0.249 

DISSOCIATION,  PER  CENT 


Pressure  in  atmospheres 


Temperature, 
degrees 
Fahrenheit 

0.1 

1.0 

10 

100 

H2O 

2,730 
3,630 
4,530 
5,430 

0.043 
1.25 

8.84 
28.4 

0.02 
0.58 
4.21 
14.4 

0.009 
0.27 
1.98 
7.04 

0.004 
0.125 
0.927 
3.33 

C02 

2,730 
3,630 
4,530 
5,430 

0.104 
4.35 
33.5 

77.1 

0.048 
2.05 

17.  6| 

54.8^ 

0.0224 
0.96 
8.63 
32.2 

0.01 
0.445 
4.09 
16.9 

Calculation  of  the  theoretical  efficiency,  taking  into  account 
both  the  variable  specific  heats  and  dissociation,  shows  that  this 
1  TIZARD  and  PYE,  The  Automobile  Engineer,  Feb.,  1921. 


ENGINE  EFFICIENCIES  AND  CAPACITIES 


21 


efficiency,  with  a  mixture  giving  maximum  efficiency,  is  repre- 
sented very  closely  by  the  equation 


E  =  1  -  (- 


0.295 


(2) 


The  heat  loss  to  the  walls  reduces  the  actual  efficiency  below  the 
theoretical  values.  With  the  very  best  design  of  cylinder  and 
optimum  operating  conditions  the  highest  attainable  indicated 
thermal  efficiency  is  given  fairly  accurately  by  the  equation 


E 


1  - 


(3) 


A  comparison  of  these  efficiency  values  is  given  in  the  following 
table.  There  are  added  the  best  results  obtained  by  Ricardo1 
on  a  special  engine  in  which  every  known  refinement  was  em- 
ployed with  a  view  to  raising  the  thermal  efficiency. 

CYCLE  AND  ENGINE  EFFICIENCIES 


Efficiency 

Compression 

ratio  r 

Air  cycle 

From  equa- 
tion (2) 

From  equa- 
tion (3) 

Ricardo's 
observed 
values 

4.0 

0.426 

0.336 

0.296 

0.277 

4.5 

0.452 

0.359 

0.314 

0.297 

5.0 

0.475 

0.378 

0.332 

0.316 

5.5 

0.494 

0.396 

0.348 

0.332 

6.0 

0.512 

0.411 

0.361 

0.346 

6.5 

0.527 

0.424 

0.375 

0.360 

7.0 

0.540 

0.437 

0.386 

0.372 

7.5 

0.553 

0.449 

0.396 

0.383 

8.0 

0.565 

0.460 

0.406 

The  difference  between  the  air-cycle  efficiency  (constant 
specific  heat  and  no  dissociation)  and  the  theoretical  efficiency 
of  the  cycle  using  imperfect  gases,  with  the  properties  given  in 
the  preceding  tables,  diminishes  as  the  explosion  temperature 
diminishes.  In  the  limiting  case  in  which  there  is  no  fuel  in  the 
charge  and  consequently  no  rise  of  temperature  at  explosion  the 
two  efficiencies  become  equal.  The  less  the  fuel  in  the  charge,  or, 

1  Proc.  Royal  Aeronautical  Society,  1920. 


22 


THE  AIRPLANE  ENGINE 


the  weaker  the  mixture,  the  more  nearly  does  the  cycle  efficiency 
approach  the  air-cycle  efficiency. 

Calculations  by  Tizard  and  Pye  show  the  cycle  efficiency  to 
vary  with  the  mixture  strength  as  in  Fig.  13.  The  curve  for 
correct  mixture  shows  the  efficiency  when  the  air-fuel  ratio  is 


258 


/1\  0. 

chemically  correct;  the  equation  to  the  curve  is  E  =  1  —  f-J 

The  20  per  cent  weak  curve  is 
calculated  for  20  per  cent  ex- 
cess of  air  which  is  usually 
about  the  limit  of  explodibility; 
the  improved  cycle  efficiency 
in  this  case  is  verified  by  en- 
gine tests  which  generally  show 
maximum  indicated  thermal 
efficiency  with  about  20  per 
cent  excess  of  air.  The  50 
per  cent  weak  curve  represents 
a  condition  which  cannot  be 
attained  in  the  normal  Otto- 
cycle  engine  as  it  gives  a  non- 
explosive  mixture;  it  can  be 
realized  by  an  injection  of  the 
fuel  into  the  compressed  air  as 
in  the  Diesel  cycle,  or  by  hav- 

-  &    gtratified    charge    ^    the 

,.     ,  .,,  ,      . 

Cylinder     With      an      explosive 

mixture  surrounding  the  ig- 
niter. In  any  case  the  reali- 
zation of  the  higher  efficiency  of  the  weak  mixture  will  be  attended 
by  reduced  engine  capacity. 

Another  point  of  importance  is  brought  out  by  the  curves  of 
Fig.  13.  The  ratio  of  cycle  efficiency  to  air-cycle  efficiency  in- 
creases with  the  ratio  of  compression;  that  is,  we  may  expect 
to  realize  a  larger  percentage  of  the  air-cycle  efficiency  as  the 
compression  ratio  increases.  The  ratio  of  the  efficiencies  for  a 
compression  ratio  of  4  is  0.685;  with  a  compression  ratio  of 
10  it  is  0.735.  This  improvement  is  shown  also  in  actual 
engines.  The  ratio  of  observed  efficiency  (Fig.  13)  to  the  air- 
cycle  efficiency  rises  from  0.65  at  a  ratio  of  compression  of  4  to 
0.685  at  a  ratio  of  compression  of  7. 


7       &       9      10 

Compression     Ratio 

FIG.    13.  —  Calculated   and   observed 
thermal     efficiencies     with    various 

strengths  of  mixture  and  compression 


ENGINE  EFFICIENCIES  AND  CAPACITIES  23 

III.  The  theoretical  indicator   diagram  is  not  realized  for 
still  another  reason.     The  admission  and  exhaust  of  the  charge 
are  attended  by  frictional  resistance  to  the  passage  of  the  gas 
through  the  carburetor,  inlet  manifold,  inlet  valve,  and  exhaust 
valve.     Moreover,  as  the  flow  of  the  gas  is  at  high  velocity,  there 
must  be  a  pressure  drop  to  bring  about  this  flow;  with  an  inlet 
velocity  of  250  ft.  per  second,  this  would  amount  to  about  0.6 
Ib.  per  square  inch.     The  frictional  resistance  and  velocity  head 
cause  a  lowering  of  the  admission  pressure,  a,  a  raising  of  the 
exhaust  pressure,  and  the  forming  of  the  "loop"   (Fig.  11)  at 
the  bottom  of  the  indicator  diagram.     This  loop  represents  the 
negative  pumping  or  fluid  friction  work  which  the  engine  has  to 
perform.     Engine  tests  indicate  that  the  pumping  work  increases 
rather  more  rapidly  than  the  square  of  the  engine  speed;  the 
actual   amount   of  the  work  depends  on  the  dimensions  and 
arrangement  of  the  engine.     At  1,000  r.p.m.  it  will  probably 
average  about  4  per  cent  of  the  indicated  work  of  an  aviation 
engine.     This   means   in   an   engine   with   120  Ib.    per  square 
inch,  brake  m.e.p.  that  the  mean  height  of  the  loop  is  5  Ib.  per 
square  inch  at  1,000  r.p.m.     The  pumping  loss  is  a  function  of 
the  gas  velocity  in  the  manifolds  and  through  the  valve  ports. 
Its  magnitude  will  vary  from  about  2  Ib.  per  square  inch  with  a  gas 
velocity  of  100  ft.  per  second  to  8  Ib.  per  square  inch  with  a  gas 
velocity  of  200  ft.  per  second.     The  indicated  work  may  properly 
be  considered  as  being  only  the  positive  loop  of  the  indicator  card; 
the  suction-exhaust  loop  is  one  of  the  engine  friction  losses. 

Minor  factors  affecting  the  area  of  the  indicator  card  are  the 
rounding  of  the  "toe"  of  the  diagram  which  results  from  the 
opening  of  the  exhaust  valve  before  the  end  of  the  expansion 
stroke  in  order  to  facilitate  exhaust,  and  the  departure  of  the 
expansion  and  compression  curves  from  the  theoretical  adiabatic 
curves. 

IV.  It  is  not  usually  practicable  to  determine  directly  the 
work  done  by  the  working  substance  in  the  engine  cylinder  or 
the  indicated  work.     In  all  tests  the  power  measured  is  the 
useful  work  or  brake  horse  power  which  is  what  remains  of  the 
indicated  work  after  some  of  it  has  been  used  up  in  overcoming 
the  friction  of  the  engine  and  in  driving  the  water  and  oil  pumps. 

Indicated  work  —  friction  work  =  useful  work 

Useful  work          , ,    ,      •    i    «•  • 
-Z—T. —  — r  =  Mechanical  efficiency 

Indicated  work 


24 


THE  AIRPLANE  ENGINE 


Friction  Losses. — The  mechanical  losses  in  an  engine  may  be 
divided  into  two  groups: 

1.  The  losses  due  to  bearing  friction  and  the  driving  of  such 
auxiliaries  as  valve  gears,  oil  and  water  pumps,  magnetos,  etc. 

2.  Piston  friction. 

Tests  by  Ricardo  show  that  to  overcome  the  first  group  a 
mean  effective  pressure  of  from  1.5  to  3  Ib.  per  square  inch  is 
usually  required — the  lowest  figure  applying  to  a  large  multi- 
cylinder  engine.  The  distribution  of  these  losses  is  about  as 
follows : 

Bearings 0 . 75  to  1 . 00  Ib.  per  square  inch 

Valve  gear 0. 75  to  0. 80  Ib.  per  square  inch 

Magnetos 0. 05  to  0. 01  Ib.  per  square  inch 

Oil  pumps 0. 15  to  0. 25  Ib.  per  square  inch 

Water  pump 0.30  to  0.50  Ib.  per  square  inch 

Total 2 . 00  to  2 . 65  Ib.  per  square  inch 

Piston  friction  is  the  largest  item  of  loss;  its  magnitude  prob- 
ably results  from  the  fact  that  the  motion  is  reciprocating  and 


1000 


1200 


1400          1600          1800          2000         2200 
Engine  Revolutions  per  Min. 

FIG.   14. — Mechanical  efficiencies  of  airplane  engines. 

that  the  film  of  oil  on  the  walls  is  more  or  less  carbonized  by  the 
high  temperatures  and  consequently  has  a  high  viscosity.  The 
magnitude  is  probably  about  7  Ib.  per  square  inch  of  piston  area. 

The  total  loss  from  bearing  friction,  piston  friction  and  fluid 
friction  in  the  best  ungeared  engines  is  from  about  10.5  to  14  Ib. 
per  square  inch  of  piston  area,  the  lower  figure  referring  to  radial 
air-cooled  engines  and  the  higher  to  water-cooled  engines. 
Taking  120  Ib.  per  square  inch  as  the  brake  m.e.p.,  these  values 
correspond  to  mechanical  efficiencies  of  92.0  and  89.5.  Tests 
(Fig.  14)  indicate  that  friction  work  increases  more  rapidly  than 
the  engine  speed  but  not  so  rapidly  as  the  square  of  the  speed. 

Taking  into  account  the  losses  enumerated,  it  is  possible  to 
arrive  at  a  fair  approximation  to  the  actual  efficiency  of  an  air- 
plane engine.  Consider  for  example  a  high-grade  engine  with 


ENGINE  EFFICIENCIES  AND  CAPACITIES  25 

a  compression  ratio  of  5.5,  using  gasoline  as  fuel.  For  every 
100  B.t.u.  (lower  heat  value)  of  heat  of  combustion  we  may 
expect  a  heat  suppression  (Item  La)  of  4  B.t.u.,  leaving  96  B.t.u. 
developed.  Of  this  quantity,  13  B.t.u.  will  go  to  the  walls  by 
radiation  and  conduction  (Item  1.6)  before  it  can  be  utilized, 
leaving  83  B.t.u.  The  theoretical  efficiency  for  a  compression 
ratio  of  5.5  is  0.396  (equation  (2)  p.  21).  The  theoretical  work 
of  the  cycle  is  0.396  X  83  =  32.9  B.t.u.  The  actual  indicated 
work  is  thus  32.9  per  cent  of  the  heat  of  perfect  combustion  of 
the  fuel  and  this  quantity  is  usually  spoken  of  as  the  indicated 
thermal  efficiency,  or  the  thermodynamic  efficiency,  Et,  of  the 
engine.  It  measures  the  efficiency  of  the  engine  in  converting 
heat  into  work. 

As  previously  stated,  this  indicated  work  is  not  readily  meas- 
urable. The  useful  or  brake  work  may  be  taken  as  85  per  cent 
(Item  IV)  of  the  indicated  work,  or  in  this  case,  0.85  X  32.9  = 
28.0  B.t.u.  The  thermal  efficiency  referred  to  b.h.p.  is  then 
28.0  per  cent.  This  quantity  may  be  compared  with  the  results 
of  tests  on  a  high-grade  engine.  Such  tests  may  be  expected 
to  show  the  consumption  of  about  0.50  Ib.  of  gasoline  per  brake 
horse  power  hour.  The  fuel  has  a  lower  heat  value  of  almost 
18,500  B.t.u.  per  pound.  The  thermal  efficiency  referred  to 

Work  of  1  b.h.p.  hour,  B.t.u.  2,545 

n  n  n   is — - 

Heat  of  combustion  of  the  fuel,  B.t.u.      0.50  X  18,500 

=  0.275,  which  agrees  very  closely  with  the  calculated  efficiency. 
A  reduction  in  any  of  the  itemized  losses  will  increase  the  final 
efficiency. 

Mean  Effective  Pressures. — The  mean  effective  pressure 
(m.e.p.)  of  a  gas  engine  is  that  gas  pressure  on  the  piston  which, 
if  maintained  constant  for  one  stroke  of  the  engine,  would  do  as 
much  work  as  is  actually  done  in  the  two  revolutions  of  the  cycle. 

In  aviation  engines  the  m.e.p.  is  practically  always  obtained 
from  the  brake  horse  power  and  is  called  the  brake  m.e.p.  It  is 
given  by  the  equation 

b.h.v.  X  33,000  b.h.p. 

brake    m.e.p.  =  • ^r—  -  =  1,083,000  ^—: — ^,  ,r  v 

Vx-X-Xn  d*XsXNXn 

4rt    X  12  X  2  X 

where  d  is  the  cylinder  diameter  in  inches,  s  is  the  stroke  in 
inches,  N  is  the  revolutions  per  minute,  and  n  is  the  number  of 
cylinders. 


26  THE  AIRPLANE  ENGINE 

The  brake  mean  effective  pressures  usually  given  are  computed 
from  the  b.h.p.;  values  range  from  70  to  135  Ib.  per  square  inch. 
The  true  m.e.p.  in  the  cylinder  is  this  value  divided  by  the  mechani- 
cal efficiency,  Em. 

Torque  and  Power. — If  p  =  actual  brake  m.e.p.,  the  average 
useful  force  exerted  in  the  cylinders  of  a  four-cycle  engine  is 

7T  1 ) 

.ndz  X  T  =  0.1964  pnd2  Ib.  This  force  is  maintained  while  the 
piston  moves  during  each  revolution  2s  in.,  or  s  -f-  6  ft.  and  the 

o 

work  done  in  foot-pounds  is  0.1964  pnd2  X  fi  =  0.03273  pnd2s. 

Torque  is  the  average  turning  moment  and  is  numerically  equal 
to  the  force  continuously  exerted  at  the  propeller  at  1  ft.  radius. 
This  is  exerted,  during  each  revolution,  over  a  distance  of  2?r  = 
6.2832  ft.  The  work  being  equal  to  that  already  computed, 
the  torque  in  pound-feet  is  Q  =  0.03273  pnd2s  -r-  6.2832  =  pnd2s 
-^  192. 

For  3  =  7,  d  =  5,  n  =  12,  p  =  120;  Q  =  1,315  Ib-ft.  The 
actual  torque  varies,  but  has  this  average  value. 

2sN 
If  S  =  piston  speed,  feet  per  minute  =  -TTT'  the  b.h.p.  is  HB  = 

*>|  Sn  -5-  33,000  =  pd2Sn  ^  168,000  =  Q  X  N  -5-  5,250.     Thus 

for  1,600  r.p.m.,  in  the  preceding  example,  HB  =  (1,315  X 
1,600)  •*-  5,250  =  401. 

Capacity  and  Volumetric  Efficiency. — The  weight  of  fuel 
mixture  taken  into  the  ideal  engine  (Fig.  10)  is  given  by  the  gas 
equation 

b  —  va) 


where  Vb  —  va  =  A  d2s  is  the  volume  of  the  mixture  admitted  and 
4 

Tm  is  the  absolute  temperature  of  the  external  air. 

Actual  engines  do  not  draw  in  weights  equal  to  that  expressed 
by  the  above  equation.  The  weight  of  mixture  actually  admitted 
is  the  difference  between  the  weight  present  at  the  points  b  and 

144  pv 
a,  Fig.  10.     The  weight  present  at  any  point  is  w  =  -fwir* 

therefore    the  'weight    admitted    is    ~j>%r°  —  —&£— 

III  b  HI  a 

The  ratio  of  the  weight  actually  admitted  to  that  which  would  be 


ENGINE  EFFICIENCIES  AND  CAPACITIES 


27 


admitted  to  the  ideal  engine  is  called  the  volumetric  efficiency,  Ev. 

<aVg\        *        144     X     14.7    (Vb    -    Vg) 

1  nn  I     *  ~D  /TT 

1  a'  K  1  m 


Writing   -  =  r,  the  above  reduces  to 


-  M 
TJ 


14.7(r  -  1)  V  Tb 

Taking  tm  =  100,  r  =  5,  pb  =  12,  tb  =  200,  pa  =  16,  ta  =  900 
we  find  Ev  =  0.76. 

The  volumetric  efficiency  is  determined  mainly  by  two  factors, 
the  temperature,  Tb,  and  the  pressure,  pb,  at  the  end  of  admission. 


,3.3 


I 


3.1 


.ci 
-[2.9 


"    200       400          600          800          1000          1200        1400         1600 
R.P.  M. 

FIG.  15. — Volumetric  efficiencies  of  hot  and  cold  engines. 


The  temperature  of  the  mixture  rises  during  admission  as  a  result 
of  the  addition  of  heat  from  the  hot  interior  surfaces  of  the  cylin- 
der. Comparative  tests  of  the  volumetric  capacity  of  an  engine 
(1)  when  being  motored  over  cold,  and  (2)  in  ordinary  operation, 
show  that  the  heating  effect  decreases  slightly  as  the  speed  of  the 
engine  increases  in  consequence  of  the  shorter  time  available  for 
the  transmission  of  heat.  The  pressure  drop,  however,  increases 
continuously  with  increase  of  engine  speed.  Figure  151  gives 
curves  of  weight  of  charge  taken  into  the  cylinder  per  revolution 
for  an  engine  of  high  valve  resistance.  The  effect  of  the  tempera- 
ture rise  in  reducing  the  volumetric  efficiency  is  from  12  to  15  per 
cent  and  would  be  appreciably  greater  but  for  the  evaporation  of 
the  fuel,  which,  through  the  abstraction  of  the  latent  heat  of  evap- 
1  JUDGE,  High-speed  Internal-combustion  Engine,  p.  161. 


28 


THE  AIRPLANE  ENGINE 


oration,  reduces  the  rise  of  temperature  of  the  charge  by  30  to 
40°F.  The  variation  of  volumetric  efficiency  of  the  Liberty-12 
with  engine  speed  is  shown  in  Fig.  16.  In  the  same  figure  there 
is  also  shown  the  volumetric  efficiency  of  the  Hispano-Suiza  300 
engine,  but  in  this  case  the  volumetric  efficiency  is  given  as  the 
ratio  of  the  weight  of  air  actually  admitted  to  the  weight  of  a 
volume  of  air  equal  to  piston  displacement  and  of  the  density 
of  the  air  in  the  inlet  manifold.  Measured  in  this  way  the  volu- 
metric efficiency  is  95  per  cent  at  1,600  r.p.m.;  this  corresponds 
to  93  per  cent  when  compared  with  air  at  room  density.  The 
broken  line  shows  the  volumetric  efficiencies  compared  with  air 
at  room  density. 


ume+ric  Efficiency,  Per  C 

^J  OO  <S  O  — 

o  p  o  o  c 

Hispctno 

-S(//y 

•3 

*^z~- 



Libe> 

TO 

.o-—11 

„ 

~fy 

5  —  -C 

'•***, 

1  —  ,  .  

? 

$ 

FIG.  16. — Volumetric  efficiencies  of  airplane  engines. 


R.RM. 


The  pressure  drop  during  admission  is  much  more  variable  in 
different  engines  than  is  the  temperature  rise.  Its  magnitude 
depends  on  the  pressure  drop  through  the  carburetor  and  the 
size  and  arrangement  of  the  manifolds  and  the  inlet  valve.  In 
every  case  it  will  increase  rapidly  with  increased  engine  speed; 
its  actual  magnitude  will  usually  be  small  for  low  speeds. 

The  pressure  drops  in  the  manifolds  of  two  engines  are  given 
in  Fig.  17.  It  will  be  seen  that  with  wide-open  throttle  the 
pressure  drop  is  nearly  proportional  to  the  engine  speed.  With  an 
engine  loaded  with  a  propeller,  change  of  speed  is  obtained  only 
by  varying  the  opening  of  the  throttle  valve;  the  manifold 
vacuum  increases  rapidly  as  the  throttle  valve  is  closed.  The 
pressure  in  the  cylinder  is  considerably  less  than  that  in  the 
manifold  because  of  the  valve  resistances. 

The  volumetric  efficiency  in  engines  of  good  design  will  be 
from  80  to  85  per  cent.  With  low  speeds  and  other  exceptionally 
favorable  conditions,  values  as  high  as  90  to  92  per  cent  have 
been  recorded. 

The    maximum    possible    volume    of    charge    admitted    per 


ENGINE  EFFICIENCIES  AND  CAPACITIES 


29 


cycle  in  the  ordinary  engine  is  the  volume  enclosed  at  the  instant 
of  valve  closure  less  the  clearance  volume.  The  admission 
valve  always  closes  past  the  dead  center.  If  the  closing  angle 
is  45  deg.  late  the  piston  will  have  returned  about  12  per  cent  of 
its  stroke  and  the  maximum  possible  volumetric  efficiency  will 
be  88  per  cent.  Occasionally  the  operating  conditions  and  induc- 
tion pipe  length  may  be  such  as  to  give  more  than  atmospheric 
pressure  in  the  cylinder  at  the  instant  of  closure  which  would 
result  in  increased  volumetric  efficiency. 


s 


\ \ 


x  j\ 


\3\ 


1200 


Hisparro  -SuigaSOOHp. 
Liberty  6 


2400 


1400          1600         1800          2000         2200 
Revolutions   per  Minu+e  of  Engine 

FIG.   17. — Intake  manifold  depressions  with  full  throttle  and  with  propeller 

load. 

It  should  be  noted  that  the  capacity  (as  affected  by  volumetric 
efficiency)  and  thermal  efficiency  of  an  engine  are  not  necessarily 
related  to  one  another.  The  diminution  in  capacity  of  an  engine 
resulting  from  heating  of  the  entering  charge,  from  a  high 
carburetor  or  inlet-valve  resistance,  or  from  diminishing  speed, 
may  or  may  not  result  in  a  change  in  efficiency,  and  the  change, 
if  it  takes  place,  may  be  either  an  increase  or  a  decrease.  A 
given  engine  may  at  one  time  be  developing  200  h.p. ;  at  another 
time  250  h.p.;  the  efficiency  may,  however,  be  the  same  in  both 
cases,  although  it  tends  to  be  lower  for  the  lower  h.p.  because  of 
the  approximate  constancy  of  engine  friction,  which  makes  the 
efficiency  referred  to  the  b.h.p.  less  at  light  loads. 

Units  of  Capacity. — In  determining  the  size  of  a  projected 
engine,  or  in  comparing  the  performance  of  existing  engines,  it 
is  desirable  to  have  some  standard  unit  for  measuring  the  specific 
capacity.  The  most  common  unit  is  the  piston  displacement  in 


30  THE  AIRPLANE  ENGINE 

cubic  inches  per  brake  horse  power,  or  the  brake  horse  power 
developed  per  cubic  foot  of  piston  displacement.  The  piston 
displacement  is  the  displacement  per  stroke  of  one  cylinder 
multiplied  by  the  number  of  cylinders.  As  the  horse  power 
varies  almost  directly  as  the  engine  speed,  the  above  units  do 
not  really  lead  to  a  satisfactory  comparison  of  engines  operat- 
ing at  different  speeds.  For  this  purpose  it  is  better  to  state 
the  capacity  at  1,000  r.p.m.,  deducing  this  capacity  from  the 
actual  performance  by  the  use  of  the  assumption  that  horse 
power  is  proportional  to  engine  speed.  For  example,  the  Lib- 
erty-12  engine,  5  by  7  in.,  develops  400  h.p.  at  1,700  r.p.m.  The 

piston  displacement  per  cylinder  is  ^  X  52  X  7  =  137.4  cu.  in. 

per  stroke;  the  total  piston  displacement  is  12  X  137.4  =  1,648.8 
cu.  in.;  the  piston  displacement  per  brake  horse  power  is  1,648.8 
-5-  400  =  4.12  cu.  in.  The  brake  horse  power  per  cubic  foot  of 
piston  displacement  is  123  -f-  4.12  =  420  h.p.;  the  piston  dis- 
placement per  brake  horse  power  at  1,000  r.p.m.  is  4.12  X 

1  700 

=  7.0  cu.  in.;  the  b.h.p.  per  cubic  foot  of  piston  displace- 


ment at  1,000  r.p.m.  is  420  X    ~     =  247.   ' 

The  Hispano-Suiza  engine,  with  718.9  cu.  in.  displacement, 
develops  150  h.p.  at  1,450  r.p.m.  This  corresponds  to  4.8  cu. 
in.  per  horse  power;  or  6.96  cu.  in.  per  horse  power  at  1,000  r.p.m. 
This  last  figure  shows  that  the  Hispano-Suiza  and  Liberty  engines 
have  practically  the  same  capacities  per  cubic  inch  of  piston 
displacement  per  minute. 

The  fixed-cylinder  radial  air-cooled  ABC  Dragonfly  nine- 
cylinder  engine,  with  1,389.3  cu.  in.  displacement,  develops  310 
h.p.  at  1,650  r.p.m.  This  corresponds  to  4.48  cu.  in.  per  horse 
power,  or  7.38  cu.  in.  per  horse  power  at  1,000  r.p.m. 

The  lower  specific  capacity  of  rotating-cylinder  engines  is 
illustrated  by  the  nine-cylinder  Gnome,  with  a  piston  displace- 
ment of  770  cu.  in.,  which  develops  104  b.h.p.  at  1,200  r.p.m. 
This  corresponds  to  7.41  cu.  in.  per  horse  power,  or  8.89  cu.  in. 
per  horse  power  at  1,000  r.p.m. 

The  above  figures  may  be  regarded  as  characteristic  of  the 
different  types. 

Tests  of  Performance.  —  The  results  obtained  on  the  test  of  an 
engine  will  vary  greatly  with  a  number  of  factors  such  as  the  air 


ENGINE  EFFICIENCIES  AND  CAPACITIES 


31 


pressure  and  temperature,  kind  of  fuel,  type  and  dimensions  of 
carburetor,  temperature  of  jacket  water  and  of  lubricating  oil, 
and  condition  of  engine.  For  example  the  Liberty  12  has  shown 
a  brake  horse  power  at  1700  r.p.m.  which  varies  from  380  to  480 
b.h.p. 

The  following  tests  are  reported  under  sea-level  conditions: 


Engine 

Revolu- 
tions 
per 
minute 

Brake 
m.e.p., 
Ib.  per 
sq.  in. 

Brake 
horse 
power 

Friction 
horse 
power 

Me- 
chanical 
efficiency 

Hispano-Suiza-180 

1,200 

114.0 

121.8 

16.8 

0.88 

1,500 
1,700 
1,900 

119.1 
117.7 
111.0 

159.0 
178.0 
187.0 

23.5 

28.4 
34.2 

0.87 
0.86 
0.84 

Liberty-12  

1,200 

118.0 

295.0 

27.6 

0.91 

1,400 
1,600 
1,800 
2,000 

119.5 
119.5 
117.6 
104.0 

348.0 
398.0 
442.0 
433.0 

38.3 
49.1 
65.4 
88.0 

0.90 
0.89 
0.87 
0.83 

A  plotting  of  test  results  on  the  Liberty  12  is  shown  in  Fig.  18. 
These  figures  illustrate  the  usual  laws  of  performance.     The 


430 
460 
440 
420 

0 

£-380 

J5360 
I 
340 

320 
300 

/ 

Me 

-A.  E 

ffici 

*ncy 

/ 

/ 

*•  —  i 

-1  —  ^ 

/ 

I  hi 

rmc 

j     f 

<ttti 

cien 

~v  , 

^•'^ 

NS 

/ 

^ 

'  ^> 

f 

/ 

~j 

f- 

A 

/ 

/ 

/ 

/ 

280 
260 

1200          1400          1600         1800         2000 
R.P.  M. 

FIG.  18. — Performance  curves  of  Liberty-12  engine. 


mean  effective  pressure,  p,  and  consequently  the  torque,  reach 
maximum  values  at  some  moderate  speed.  The  power  increases 
with  increasing  speed,  but  at  a  rate  which  diminishes  after  the 


32  THE  AIRPLANE  ENGINE 

maximum  value  of  p  has  been  reached.  If  p  were  constant  the 
power  would  vary  directly  with  the  speed.  Cylinder  cooling 
reduces  p  at  low  speeds;  high  resistance  through  ports  and 
passages  reduces  volumetric  efficiency  and  p  at  high  speeds  (Fig. 
16).  Maximum  power  is  reached  when  the  rate  of  decrease  of  p 
with  engine  speed  is  equal  to  the  rate  of  increase  of  engine  speed. 

Figure  79  gives  results  of  trials  on  a  230  h.p.,  six-cylinder  Benz 
engine.  Here  the  failure  of  the  power  to  increase  proportionately 
to  speed  is  clearly  shown.  Maximum  mean  effective  pressure 
occurs  at  1,050  r.p.m.  and  maximum  power  at  1,650  r.p.m.  The 
fuel  consumption  rate  is  also  shown.  The  economy  is  practically 
constant  over  the  speed  range  900  to  1,200  r.p.m. 

It  will  be  observed  that  throttling  the  engine  increases  the 
fuel  consumption  per  horse  power.  The  full-line  curves  show 
the  performance  when  the  throttle  is  wide  open  and  the  engine 
is  loaded  until  it  assumes  the  desired  speed.  The  broken  lines 
show  the  performance  when  the  engine  had  the  propeller  load 
only;  in  this  case  the  engine  speed  can  be  reduced  below  1,400 
r.p.m.  only  by  partly  closing  the  throttle;  speeds  above  1,400 
r.p.m.  are  not  possible  with  the  propeller  load. 

Figure  143  is  for  a  rotary-cylinder  engine.  In  this  case  the 
effective  horse  power  is  the  indicated  horse  power  minus  the  engine 
friction  loss  and  minus  the  windage  loss;  the  rapid  increase  of 
windage  loss  when  the  engine  speed  increases  makes  the  net 
horse  power  a  maximum  at  about  1,250  r.p.m. 

Correction  to  Standard  Atmospheric  Conditions. — The  pub- 
lished results  of  engine  tests  may  give  either  the  actual  horse 
powers  observed  or  these  horse  powers  corrected  to  some  standard 
atmospheric  condition.  The  latter  is  much  preferable  as  it  will 
permit  an  immediate  comparison  of  engine  performances.  The 
best  measure  of  capacity  is  the  brake  m.e.p.  That  engine  which 
has  the  maximum  m.e.p.  is  developing  a  horse  power  'on  the 
smallest  piston  displacement.  But  to  make  such  direct  com- 
parison the  operating  conditions  must  be  the  same  or  else  the 
results  must  be  corrected  to  allow  for  differences.  Such  correc- 
tions can  be  readily  applied  to  differences  in  atmospheric  pressure 
and  temperature.  It  has  been  commonly  assumed  that  the 
horse  power  developed  is  proportional  to  the  density  of  the 
atmosphere.  The  density  is  proportional  to  the  atmospheric 
pressure,  and  inversely  proportional  to  the  absolute  temperature. 
If  the  standard  conditions  are  14.7  Ib.  pressure  (29.92  in.  of 


ENGINE  EFFICIENCIES  AND  CAPACITIES 


33 


mercury)   and  32°F.,   and  the  observed  horse  power   is  P   at 
pressure  p  and  temperature   t,   the   corrected  horse  power  is 

14.7     460  +  t 


PC=P 


P 


492 


In  most  of  the  published  tests  the  correction  to  standard 
conditions  has  been  made  by  use  of  this  equation.  Tests  at  the 
Bureau  of  Standards  indicate,  however,  that  the  temperature 
correction  in  this  equation  is  excessive  and  that  more  accurate 
results  are  obtained  from  the  equation 

14.7      920-M 

p  952 

The  m.e.p.  is  corrected  in  exactly  the  same  manner. 

There  is  no  method  for  directly  comparing  two  engines  which 
are  using  different  fuels  or  mixtures  of  different  strengths 

Influence  of  Strength  of  Mixture  on  Capacity  and  Efficiency. — 
The  effect  of  strength  of  mixture  has  been  investigated  by 
Berry1  on  automobile  engines;  his  results  are  supported  by 
the  investigation  of  others  on 
both  automobile  and  aviation 
engines  (see  p.  260).  For  any 
constant  engine  speed  and  con- 
stant throttle  opening,  they 
show  (Fig.  19)  that  the  max- 
imum power  is  obtained  with  a 
comparatively  rich  mixture, 
and  that  for  maximum  effici- 
ency a  weaker  mixture  must 
be  used.  As  the  throttle  is 
closed  the  mixture  for  maxi- 
mum efficiency  (Fig.  21)  be- 
comes richer  and  at  the  lowest 
loads  coincides  with  that  for 
maximum  power.  The  speed 
of  the  engine  has  no  appreci- 
able influence  on  the  variation  of  engine  power  with  strength  of 
mixture  (Fig.  22).  Maximum  power  (Fig.  20)  is  obtained  with 
0.08  Ib.  of  gasoline  per  pound  of  air,  or  12^  lb.  of  air  per  pound 
of  gasoline.  Maximum  efficiency  is  obtained  with  a  mixture 
of  15  to  16  lb.  of  air  per  pound  of  gasoline  at  full  load,  but  this 

1  Trans.  Am.  Soc.  M^Ji.  Eng.,  1919. 
3 


0.07    0.08     0.09     0.10     O.I  I      0.12    0.13 

Pounds  of  Gasoline  per  Pound  of  Air  in  Mixture 

FIG.  19.  —  Variation  of  power  and 
thermal  efficiency  with  strength  of  mix- 
ture, at  full  throttle. 


34 


AIRPLANE  ENGINE 


mixture  must  be  made  richer  as  the  load  diminishes  and  becomes 
Ib.  of  air  per  pound  of  gasoline  at  lowest  loads. 


0.05  0.06  0.07  0.08   009  0.10  0.11   0.12  0.13  0.05  0.06  0.01  0.08  0.09  0.10  0.11    0.2   0.13 

Pounds  of  Gasoline   per  Pound  of  Air  in   Mixture 

FIG.  20.  FIG.  21. 

FIG.  20. — Variation  of  engine  power  with  strength  of  mixture  at  constant  engine 

torque  and  varying  speed. 

FIG.  21. — Variation  of  thermal  efficiency  with  strength  of  mixture  at  constant 
engine  torque  and  varying  speed. 


0.05  0£6  0.07  tt08  0.09  0.10  0.11  0.12  0.13 
founds  of  Gasoline  per  Fbund  of  Air  in  Mixture 


20 


14          12  10  8 

Ratio  of  Air  to  Fuel    by  Weight 

FIG.  23. 


FIG.  22. 
FIG.  22.— Variation  of  engine  power  with  strength  of  mixture  at  constant  engine 

speed  and  varying  torque. 

FIG.  23. — Maximum  thermal  efficiencies  of  certain  fuels  with  varying  strength 

of  mixture. 

Tests  by  Watson1  on  an  automobile  engine   with   gasoline, 
benzol,  and  wood  alcohol  as  fuel  show  (Fig.  23)  the  variation  in 
1  Proc.  Inst.  Aut.  Eng.,  1914. 


ENGINE  EFFICIENCIES  AND  CAPACITIES 


35 


efficiency  of  these  fuels  with  strength  of  mixture.  The  com- 
parison given  by  these  curves  is  not,  however,  complete,  since 
the  same  compression  ratio  was  used  for  all  three  fuels.  With 
alcohol  it  is  possible  to  increase  the  compression  pressure  con- 
siderably without  danger  of  preignition  and  without  producing 
excessive  explosive  pressures;  with  benzol  the  compression  ratio 
can  similarly  be  increased  slightly.  The  efficiency  with  alcohol 
could  probably  be  raised  to  at  least  35  per  cent  by  the  use  of  a 
higher  compression  ratio. 

Influence  of  Air  Temperature  on  Capacity  and  Efficiency. — 
As  already  pointed  out  in  the  discussion  of  volumetric  efficiency 
(p.  27)  the  temperature  of  the  air  admitted  to  the  cylinder  has 


005  0.06  0.07  0.08  OX)9  0.10  OJ!   0.12  0.13 


05  0.06  0.07  0.08  0.09  0.10  0.11   0.12  0.13 
Pounds  of  Gasolene  per  Pound  of  Air  in  Mixture  • 

FIG.  24.  FIG.  25. 

FIG.  24. — Variation  of  engine  power  with  strength  of  mixture  at  various  air 

temperatures  and  full  throttle. 
FIG.  25. — Variation  of  thermal  efficiency  with  strength  of  mixture  at  various 

air  temperatures  and  full  throttle. 

a  considerable  influence  on  the  power  developed.  The  tests 
at  the  Bureau  of  Standards  show  the  diminution  in  power  with 
increase  in  temperature  to  be  proportional  to  half  the  increase 
in  absolute  temperatures,  for  the  range  of  temperature  from 
4  to  120°F.  For  example,  if  the  absolute  temperature  of  the 
air  increases  from  500  to  580°,  or  16  per  cent,  the  power  of  the 
engine  will  decrease  8  per  cent.  Berry's  tests  on  automobile 
engines  (Figs.  24  and  25)  show  that  this  law  does  not  hold  for  a 
higher  temperature  range.  The  decrease  in  power  is  roughly 
proportional  to  one-third  the  increase  in  temperature  (Fig.  24). 
The  efficiency  (Fig.  25)  is  also  seen  to  diminish  with  increase  of 
air  temperature  but  through  a  much  smaller  range. 

Tests  by  Berry  with  air  at  a  temperature  lower  than  80°F. 


36 


THE  AIRPLANE  ENGINE 


showed  a  rapid  falling  off  in  capacity  and  efficiency.  These 
tests,  however,  were  carried  out  with  a  commercial  gasoline 
of  low  volatility  as  compared  with  the  gasolines  specified  for 
airplane  engines.  In  all  cases  maximum  power  is  obtained  with 
the  lowest  air  temperature  which  will  permit  satisfactory  dis- 
tribution and  vaporization  of  the  fuel.  This  temperature 
depends  not  only  on  the  volatility  of  the  fuel  but  also  upon  the 
manifold  design.  A"  hot  spot "  between  the  carburetor  and  mani- 
fold (heated  by  the  exhaust  gases),  on  which  the  liquid  spray 
from  the  carburetor  impinges,  causes  vaporization  of  part  of  the 
fuel  without  heating  up  the  air  appreciably  and  is  found  to 
result  in  a  better  distribution  of  the  mixture  to  the  different 
cylinders  and  in  improved  engine  operation  when  the  air  supply 
is  cold.  With  this  device  it  is  possible  to  lower  the  temperature 
range  of  the  entering  air  somewhat  without  a  falling  off  in  capacity 
or  efficiency. 

Influence  of  Throttling  on  Efficiency. — It  is  generally  found 
that  thermal  efficiency  tends  to  increase  as  the  power  is  cut  down 


0.65 
J 

7:  0.60 


0.55 


0.50 


0.45 


half 


oad 


"800  1000  1200  1400  1600  1800 

FIG.  26. — Variation  of  specific  fuel  consumption  with  engine  speed  at  various 

throttle  positions. 

from  maximum  to  three-fourths  load.  This  phenomenon  is  prob- 
ably due  mainly  to  improvement  in  the  mixture  at  partial  load. 
The  carburetor  is  set  to  give  maximum  power  for  full  throttle, 
which,  as  just  shown,  is  obtained  with  an  over-rich  and  conse- 
quently inefficient  mixture.  If  the  mixture  becomes  leaner  at 
partial  throttle,  the  economy  will  improve.  What  actually 
happens  will  depend  primarily  on  the  characteristics  of  the 
carburetor  used. 


ENGINE  EFFICIENCIES  AND   CAPACITIES 


37 


In  Fig.  26  are  given  the  fuel  consumptions  per  brake-horse- 
power hour  of  a  Liberty  12,  at  full,  three-fourths  and  one-half 
loads.  It  will  be  seen  that  the  efficiencies  at  full  and  three- 
fourths  loads  are  substantially  the  same,  but  that  there  is  a 
marked  falling  off  at  half  load.  The  efficiency  is  again  seen  to 
increase  with  engine  speed. 

Influence  of  Compression  Ratio  on  Capacity. — Tests  of  a  150- 
h.p.  Hispano-Suiza  engine  at  1,500  r.p.m.  with  various  compres- 
sion ratios  show  the  maximum  attainable  brake  horse  power  to 
have  been  as  follows: 

Ratio  of  compression ,  .  4.7  5.3          6.2 

Maximum  brake  horse  power .  .        160 . 0  165 . 0       169 . 0 

The  percentage  increase  of  power  with  increase  of  compression 
ratio  is  about  the  same  as  in  the  tests  of  a  Liberty  12  engine, 
which  at  1,600  r.p.m.  gives  the  following  results: 

Ratio  of  compression 4.9          5.5 

%      Maximum  brake  horse  power 380.0      398.0 

The  Influence  of  Revolutions  per  Minute  on  Capacity  is  shown 
in  all  performance  curves  (see  Figs.  48,  50,  53,  etc.).  The  fall  in 


1000       1200         1400        1600         1800       2000       2200      2400 
Revolutions  per  Minute  of  Engine 

FIG.  27. — Variation  of  mean  effective  pressure  with  engine  speed  for  airplane 

engines. 

brake  m.e.p.  (Fig.  27)  causes  the  b.h.p.  to  go  through  a  maxi- 
mum; the  efficiency  practically  always  increases  with  speed,  but 
becomes  a  maximum  before  the  engine  reaches  maximum  horse 
power. 

Influence  of  Jacket-water  Temperature  on  Capacity. — Figure 
28  shows  the  effect  on  the  capacity  of  a  Liberty  12  engine  of 
varying  the  temperature  of  the  jacket  water.  The  amount  of 
water  circulated  was  constant  at  any  given  speed  but  the  inlet 
temperature  was  varied,  thereby  giving  the  series  of  outlet 


38 


THE  AIRPLANE  ENGINE 


temperatures  indicated.  It  will  be  seen  that  the  power  increases 
as  the  cooling-water  temperature  decreases  to  about  100°F.  At 
200°F.  and  1,800  r.p.m.  the  power  is  only  417  h.p.  while  at  90° 
it  is  436  h.p. 


420 

400 
u 
|380 

(X 

4> 

<o 

o360 

• 
*340 
u 

CO 

320 
300 

280 

Correci-ec 

1h29 

92Inc 

hes  of  Mercury   ^ 

'/'t? 

2 

^/ 

m 

/ 

A 

'// 

1m 

A 

w- 

'200  °F. 

M 

*s 

I50°F. 
I30°F. 
JIO"f 

/f 

'///f 

^    X 

//< 

2 

^~ 

-90°f 

///// 

2 

/// 

///// 

w 

y 

0         II          12        13         14        15         16        17        18         19 

R.  P.  M.   Hundreds 

FIG.  28. — Variation  of  engine  power  with  speed  for  various  jacket-water  dis- 
charge temperatures. 

Various  tests  have  demonstrated  that  the  friction  horse  power 
decreases  with  increase  of  jacket-water  temperature,  so  that  the 
above  increase  of  brake  horse  power  with  lower  jacket- water 
temperature  must  have  resulted  from  a  still  greater  increase  in  in- 


3 

c 
|30 

JlO 

i_ 

n^ 

1 

haust  .- 

ou 
40 
30 
20 

Residual 

B. 

H.P. 

Jack 

\—£- 

1200 


1400 


1SOO 
R.  R  M. 


1800 


ZOOO 


JPIG>  29. — Typical  heat  balance  of  airplane  engine. 

dicated  horse  power.  The  increase  is  primarily  due  to  improve- 
ment in  volumetric  efficiency  resulting  from  less  heating  up  of  the 
charge  as  it  enters.  This  phenomenon  is  shown  in  Fig.  15. 

Heat  Balance. — Of  the  total  heat  of  combustion  of  the  fuel 
admitted  to  the  engine  cylinder,  part  is  converted  into  brake 


ENGINE  EFFICIENCIES  AND  CAPACITIES  39 

horse  power,  part  goes  to  the  water  jacket,  part  escapes  as  heat 
in  the  exhaust  gases,  and  the  rest  is  lost  in  various  ways  such  as 
incomplete  combustion  and  radiation  and  conduction  from  the 
engine.  A  typical  heat  balance  for  the  Liberty  12  is  shown  in 
Fig.  29  in  which  the  percentages  are  in  terms  of  the  higher  heat 
value  of  the  fuel.  It  will  be  observed  that  the  heat  distribution 
does  not  change  notably  with  engine  speed. 


CHAPTER  III 
ENGINE  DYNAMICS 

Turning  Moment. — The  pressure  on  any  single  piston  of  a 
four-cycle  engine  is  varying  continuously  throughout  the  cycle. 


t>vv 

£  40° 
g 

\ 

k 

\ 

^300 

JS 

_J 

«     onn 

\ 

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1_     CV.(J 

*•    100 
0 

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•  

•  

'  

—  s 

•—  ~~  — 

—    — 

= 

== 

—  - 

—  — 

*ai 

0 

234567 

Piston  Travel,  In. 
FIG.  30. — Indicator  card  of  the  Liberty-12  engine. 

If  the  indicator  card  is  as  in  Fig.  30,  the  resulting  total  gas  pressure 
on  the  piston  of  a  5-in.  diameter,  7-in.  stroke  engine  for  various 


5E 

A-TohtI6a>. 

iPressu 

re  on  Piston 

7000 

i 
y 

B-  Inertia  Forces  of  1700  Rpm. 

£-  Resultant  Pressure  along  Cy  It 

Hkfi 

x/s 

.     5000 
3    4000 
«T    3000 

§2000 

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u 

p\ 

•^  —  . 

^ 

A> 

£     1000 
*          0 

-1000 
-2000 

V 

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( 

^» 

^ 

/ 

/ 

^ 

'S 

/ 

1 

? 

B 

>/ 

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/ 

X 

^r 

/ 

s 

V 

/ 

r 

B>\ 

^i 

* 

Crank  Angle  Decjrees 
FIG.  31. — Forces  acting  on  the  crank  pin  of  the  Liberty-12  engine. 

successive  crank  positions  is  represented  by  the  curve  A,  Fig. 
31.  The  pressure  transmitted  to  the  crankpin  is  modified, 
however,  by  the  inertia  of  the  reciprocating  masses  of  the  piston 

40 


ENGINE  DYNAMICS  41 

and  connecting  rod.  During  the  first  part  of  each  stroke  these 
masses  are  being  accelerated;  during  the  second  part  they  are 
retarded.  Hence  the  net  useful  force  acting  on  the  crankpin 
in  a  direction  parallel  to  the  cylinder  axis  is  alternately  less  or 
greater  than  that  shown  by  the  curve  A  . 

Let  W  =  Weight   of   reciprocating   parts,    pounds   (complete 

piston  and  half  of  rod). 
n  =  Revo'utions  per  minute. 
a  =  Angle  turned  through  by  crank,  starting  from  its 

uppermost  position,  degrees. 

r  =  Crank  radius,  feet  (half  the  stroke  of  the  engine). 
I  =  Length  of  connecting  rod  (center  to  center  of  pins), 

feet. 

Then  the  accelerating  force  at  any  moment,  in  pounds,  is 
Pa  =  0.00034  Wn2r(cos  a  ±     cos    2a),   approximately,   the  + 


sign  being  used  for  the  down  stroke  and  the    —  sign  for  the  up 

7* 

-^cos2a)  i 


7* 

stroke.     The  quantity  (cos  a  ±  -cos2a)  is  an  approximation;  a 


2a  +  sin  4a 
more  correct  expression  is  cos  a  ±  -  —  Values  of 


this  quantity  for  the  range  of  ratios  of  I  to  r  common  in  airplane 
engines  are  given  in  Table  1.  The  minus  sign  indicates  negative 
acceleration  from  0  deg.  to  180  deg.,  and  positive  acceleration 
from  180  deg.  to  360  deg.  Calculations  made  for  the  Liberty-12 
engine  give  the  results  shown  in  Fig.  31.  The  indicator  card 
(Fig.  30)  is  plotted  for  18  per  cent  clearance  and  a  brake  m.e.p. 
of  123  Ib.  per  square  inch,  the  exponents  of  the  compression 
and  expansion  curves  being  taken  as  1.32.  The  cylinder  is  5 
by  7  in.,  the  connecting  rod  12  in.  long  and  the  weights  of 
reciprocating  parts  are:  piston  complete  with  pin,  4.838  Ib.; 
upper  half  of  connecting  rod,  1.225  Ib.;  total,  6.063  Ib.  The 
engine  is  assumed  to  make  1,700  r.p.m.  The  inertia  forces.  P«. 
calculated  from  the  preceding  equation,  are  plotted  as  curve 
B,  Fig.  31.  The  algebraic  sum  of  gas  pressures  A  and  inertia 
pressures  B  is  shown  by  the  resultant  pressure  curve  C. 


42 


THE  AIRPLANE  ENGINE 


TABLE  1. — INERTIA  FACTORS 


-^cos2a  +  sin4  a 


cos  a  ± 


Crank 
angle, 
degrees 

/ 
=  4 
r 

'--3.75 

r 

'-=3.5 
r 

'-  =  3.25 
r 

-  =  3.0 

r 

Crank 
angle, 
degrees 

0 

.2500 

.2667 

.2857 

.3077 

1  .  3333 

360 

5 

.2426 

.2590 

.2778 

.2995 

1  .  3249 

355 

10 

.2204 

.2362 

.2543 

.2752 

1  .  2997 

350 

15 

.1839 

.1986 

.2155 

.2351 

1  .  2580 

345 

20 

.1335 

.1468 

.1621 

.1798 

1  .  2006  - 

340 

25 

.0702 

.0817 

.0948 

.1102 

1.1283 

335 

30 

0.9950 

1.0042 

1.0149 

1  .  0274 

1  .  0423 

330 

35 

0.9091 

0.9158 

0.9236 

0.9328 

0.9440 

325 

40 

0.8140 

0.8179 

0.8225 

0.8281 

0.8349 

320 

45 

0.7112 

0.7121 

0.7133 

0.7149 

0.7172 

315 

50 

0.6026 

0.6004 

0.5980 

0.5955 

0.5929 

310 

55 

0.4899 

0.4846 

0.4787 

0.4719 

0.4643 

305 

60 

0.3751 

0.3668 

0.3573 

0.3465 

0.3338 

300 

65 

0.2601 

0.2490 

0.2363 

0.2215 

0.2041 

295 

70 

0,1468 

0.1332 

0.1175 

0.0992 

0.0776 

290 

75 

0.0368 

0.0211 

0.0030 

-0.0182 

-0.0434 

285 

80 

-0.0682 

-0.0854 

-0.1055 

-0.1288 

-0.1567 

280 

85 

-0.1669 

-0.1851 

-0.2062 

-0.2309 

-0.2605 

275 

90 

-0.2582 

-0.2767 

-0.2981 

-0.3234 

-0.3536 

270 

95 

-0.3412 

-0.3594 

-0.3805 

-0.4052 

-0.4348 

265 

100 

-0.4155 

-0.4327 

-0.4528 

-0.4761 

-0.5040 

260 

105 

-0.4809 

-0.4965 

-0.5146 

-0.5358 

-0.5610 

255 

110 

-0.5373 

-0.5509 

-0.5665 

-0.5848 

-0.6064 

250 

115 

-0.5851 

-0.5962 

-0.6090 

-0.6237 

-0.6411 

245 

120 

-0.6249 

-0.6332 

-0.6427 

-0.6535 

-0.6662 

240 

125 

-0.6573 

-0.6625 

-0.6685 

-0.6752 

-0.6829 

235 

130 

-0.6830 

-0.6852 

-0.6875 

-0.6901 

-0.6927 

230 

[135 

-0.7030 

-0.7021 

-0.7009 

-0.6993 

-0.6970 

225 

140 

-0.7181 

-0.7142 

-0.7096 

-0.7040 

-0.6972 

220 

145 

-0.7292 

-0.7225 

-0.7137 

-0.7055 

-0.6944 

215 

150 

-0.7370 

-0.7279 

-0.7172 

-0.7047 

-0.6898 

210 

155 

-0.7423 

-0.7310 

-0.7178 

-0.7025 

-0.6843 

205 

160 

-0.7459 

-0.7326 

-0.7173 

-0.6997 

-0.6788 

200 

165 

-0.7480 

-0.7333 

-0.7163 

-0.6968 

-0.6738 

195 

170 

-0.7492 

-0.7334 

-0.7153 

-0.6944 

-0.6700 

190 

175 

-0.7498 

-0.7334 

-0.7146 

l  -0.6929 

-0.6675 

185 

180 

-0.7500 

-0.7333 

-0.7143 

-0.6923 

-0.6667 

180 

ENGINE  DYNAMICS 


43 


The  resultant  pressure,  P,  curve  C,  acts  along  the  axis  of  the 
cylinder.  The  force  acting  along  the  connecting  rod,  Fig.  32, 
is 

PE  =  P  +  COS  6. 

The  component  acting  tangentially  to  the 
crankpin  circle  is 

PQ  =  PE  sin  (a  +  b)  =  P  sec  b  sin  (a  +  b)- 

Table  2  gives  values  of  the  tangential 
factor  [sec  b  sin  (a  +  &)]• 

The  angles  a  and  6  are  connected  by  the 
equation 

sin  b  =  j  sin  a. 

Consequently, 

n   .        /., 
PO  =  P  sin  a  (1  H 

The  torque  or  turning  moment  applied 
to  the  crank  at  any  crank  angle  a  is 

T  =  PQr 

Figure  33  shows  the  torque  variation  for  a 
single    cylinder  of  the  Liberty   12   engine.       FIG.   32.— Diagram 
Since  the  brake  m.e.p.  is  123  lb.,  the  engine    showing  effect  of  obiiq- 

F  uity  of  connecting  rod. 

horse  power  per  cylinder  is 


123  X  ^2  X  (~  X  52)    X  850 


33,000 


=  36.3 


The  mean  torque  at  the  propeller  must  lead  to  the  same  result: 
2irnT  =  36.3  X  33,000,  or,  T  =  112;  that  is,  the  mean  torque 
per  cylinder  is  112  Ib.-ft.  The  mean  torque  at  the  crankpin 
as  determined  from  Fig.  33  is  greater  than  this  by  the  torque 
required  to  overcome  the  frictional  resistance  of  one  cylinder, 
or  one-twelfth  of  the  total  frictional  torque  of  the  engine.  If 
the  mechanical  efficiency  of  the  engine  is  85  per  cent,  the  indicated 
mean  effective  pressure  is  lo%&  X  123  lb.  per  square  inch  and 
the  mean  torque  per  cylinder  is  10%5  X  112  Ib.-ft.  The  total 
horse  power  for  the  12-cylinder  engine  is  12  X  36.3  =  436  and 
the  mean  total  crankshaft  torque  is  12  X  112  =  1,345  Ib.-ft. 


44 


THE  AIRPLANE  ENGINE 

TABLE  2. — TANGENTIAL  FACTORS 

sin  (a  +  b) 
cos  b 


Crank 
angle, 
degrees 

I 
-  =  4 

r 

l-  =  3.75 
r 

i-j 

-  =  3.25 
r 

L-«. 

r 

Crank 
angle, 
degrees 

0 

0.0000 

0.0000 

0.0000 

0.0000 

0.0000 

360 

5 

0.1089 

0.1103 

0.1119 

0.1139 

0.1161 

355 

10 

0.2164 

0.2193 

0.2226 

0.2264 

0.2307 

350 

15 

0.3214 

0.3257 

0.3305 

0.3360 

0.3425 

345 

20 

0.4227 

0.4281 

0.4343 

0.4415 

0.4499 

340 

25 

0.5189 

0.5254 

0.5329 

0.5415 

0.5515 

335 

30 

0.6091 

0.6165 

0.6250 

0.6314 

0.6464 

330 

35 

0.6923 

0.7003 

0.7098 

0.7206 

0.7333 

325 

40 

0.7675 

0.7761 

0.7860 

0.7974 

0.8108 

320 

45 

0.8340 

0.8429 

0.8529 

0.8647 

0.8786 

315 

50 

0.8914 

0.9001 

0.9101 

0.9219 

0.9358 

310 

55 

0.9391 

0.9475 

0.9572 

0.9685 

0.9819 

305 

60 

0.9770 

0.9847 

0.9938 

.0041 

1.0167 

300 

65 

.0046 

.0116 

1.0195 

.0290 

1.0401 

295 

70 

.0223 

.0282 

1  .  0352 

.0430 

1.0524 

290 

75 

.0303 

.0349 

1.0401 

.0464 

1.0539 

285 

80 

.0290 

.0320 

1.0356 

.0399 

1.0452 

280 

85 

.0186 

.0202 

1.0221 

.0242 

1  .  0268 

275 

90 

.0000 

.0000 

1.0000 

1.0000 

1.0000 

270 

95 

0.9739 

0.9723 

0.9703 

0.9680 

0.9656 

265 

100 

0.9408 

0.9376 

0.9339 

0.9296 

0.9245 

260 

105 

0.9016 

0.8970 

0.8916 

0.8853 

0.8780 

255 

110 

0.8570 

0.8511 

0.8443 

0.8364 

0.8268 

250 

115 

0.8082 

0.8011 

0.7930 

0.7836 

0.7723 

245 

120 

0.7551 

0.7473 

0.7384 

0.7278 

0.7153 

240 

125 

0.6989 

0.6907 

0.6811 

0.6697 

0.6563 

235 

130 

0.6406 

0.6320 

0.6219 

0.6102 

0.5962 

230 

135 

0.5801 

0.5713 

0.5613 

0.5495 

0.5356 

225 

140 

0.5181 

0.5094 

0.4997 

0.4882 

0.4748 

220 

145 

0.4549 

0.4468 

0.4375 

0.4267 

0.4140 

215 

150 

0.3908 

0.3835 

0.3750 

0.3652 

0.3536 

210 

155 

0.3263 

0.3198 

0.3124 

0.3038 

0.2936 

205 

160 

0.2614 

0.2559 

0.2498 

0.2425 

0.2339 

200 

165 

0.1962 

0.1920 

0.1872 

0.1817 

0.1751 

195 

170 

0.1309 

0.1280 

0.1247 

0.1209 

0.1166 

190 

175 

0.0654 

0.0640 

0.0624 

0.0604 

0.0582 

185 

180 

0.0000 

0.0000 

0.0000 

0.0000 

0.0000 

180 

With  the  firing  order  used  on  the  Liberty  engine  and  with 
a  Vee  angle  of  45  deg.,  the  firing  intervals  between  two  cylinders 
on  any  one  crank  are  315  and  405  deg.  of  crank  revolution.  The 
turning  moment  on  any  one  crank  is  obtained  by  superimposing 


ENGINE  DYNAMICS 


45 


IUW 

\ 

\ 

oUU 

£ 

\ 

\ 

5  60° 
-»- 

|   400 
c^?00 

*  . 

-200 
-400( 

FIG.  33.—  Tu 

\ 

\ 

\ 

r, 

\ 

a 

r\ 

Me 

an 

\Jo 

rqu 

•>  i 

\ 

\ 

\ 

\ 

r\ 

\ 

\ 

1?0 

\ 

/ 

\ 

\ 

I 

^ 

? 

\ 

1 

\ 

J 

\ 

J 

)           100         200        300       400        500       600        700 
Crank  Angle,  Degrees 

riling  moment  for  a  single  cylinder  of  the  Liberty-12  engine 

100         200        300        400         500        600        700 

Crank  Angle, Degrees 
FIG.  34. — Turning  moment  on  each  crank  of  the  Liberty-12  engine. 


46 


THE  AIRPLANE  ENGINE 


two  torque  curves  like  those  of  Fig.  33  with  a  phase  lag  of  315 
deg.  Adding  the  ordinates  of  two  such  curves,  as  in  Fig.  34, 
gives  the  total  torque  on  one  crank. 

The  six  cranks  of  this  engine  are  spaced  at  angular  intervals 
of  120  deg.  The  total  torque  on  the  crankshaft  is  obtained  by 
superimposing  six  curves  like  the  resultant  curve  of  Fig.  34,  with 
angular  intervals  of  120  deg.,  and  taking  the  algebraic  sums  of 
ordinates  at  the  various  crank  positions.  This  process  gives  the 
curve  of  Fig.  35.  The  torques  and  torque  ratios  are  as  follows: 


ONE  ONE 

CYLINDER      CRANK 


WHOLE 

ENGINE 


Maximum  crankshaft  torque,  pound-feet 1 , 030     1 , 240     1 , 670 

Ratio  of  maximum  to  mean  torque 9.2        5 . 54       1 . 24 


100 


200 


600 


700 


300         400        500 
Crank  Angle,  Degrees 
FIG.  35. — Torque  at  propeller  end  of  the  crankshaft  of  the  Liberty-12  engine. 

The  ratio  of  maximum  to  mean  torque  varies  with  the  angle  of 
the  Vee.  For  5-  by  7-in.  cylinders  at  120  Ib.  mean  effective 
pressure,  the  following  ratios  hold  for  torques  on  one  crank: 


Ts. 


45  DEO. 

5.2 

-2.7 
7.9 


60  DEG. 
5.1 

-2.3 

7.4 


75  DEG. 

5.6 

-1.7 

7.4 


90  DEO. 

4.8 

-1.5 

6.3 


Here 


=  maximum  torque  -r-  mean  torque, 
Tz  =  minimum  torque  -f-  mean  torque, 
TZ  =  range  of  torque  -f-  mean  torque. 
For  the  whole  engine,  ratios  are  as  follows: 


8-cylinder. 


12-cylinder. 


VEE  ANGLE 
90 
75 
60 
45 
60 
45 


Ti 

1.40 
1.42 
1.70 


2.14 
1.13 
1.25 


0.66 

0.18 

-0.13 

-0.26 

0.86 

0.89 


ENGINE  DYNAMICS 


47 


In  both  tables  negative  signs  indicate  reversal  of  direction  of  tor- 
que. The  minimum  torque  variation  is  seen  to  occur  with  equal 
firing  intervals  (eight-cylinder,  90-deg.,  and  12-cylinder,  60-deg. 
Vee  engines).  The  great  increase  in  this  variation  as  the  Vee 
angle  of  the  eight-cylinder  engine  is  diminished  is  very  marked. 
A  smooth  curve  of  crankshaft  turning  moment,  approximating 
as  closely  as  possible  to  the  mean  torque  line,  is  in  every  way 
desirable.  This  can  best  be  obtained  by  the  use  of  a  plurality 
of  cylinders  with  equal  firing  intervals.  The  greater  the  number 
of  cylinders,  the  more  uniform  is  the  torque.  The  firing  order 
of  the  cylinders  is  unimportant  from  this  standpoint,  as  long  as 
the  firing  interval  is  constant.  The  firing  order  is  of  the  utmost 
importance,  however,  in  relation  to  the  balancing  of  forces  and 
the  stresses  in  the  engine.  Smoothness  of  running  depends  on 
the  magnitude  of  the  areas  enclosed  between  the  total  torque 
curve  and  the  mean  torque  line  (Fig.  35).  Areas  above  the  line 
represent  work  done  by  the  engine  in  excess  of  the  resisting  pro- 
peller torque  and  lead  to  acceleration;  areas  below  the  line 
represent  a  deficiency  in  engine  work  and  a  consequent  slowing 
down.  In  ordinary  engine  practice  a  flywheel  is  used  to  absorb 
the  excess  and  make  up  the  deficiency  without  permitting 
excessive  change  in  engine  speed.  In  an  airplane  engine  the 
propeller  takes  the  place  of  a  flywheel;  its  large  radius  of  gyration 
enables  it  to  absorb  a  considerable  amount  of  excess  work  with 
only  a  very  small  increase  in  speed  of  rotation.  Moreover,  the 
resisting  torque  varies  as  the  square  of  the  angular  velocity  and 
hence  increases  notably  with  small  increases  in  rotative  speed. 
The  influence  of  the  number  of  cylinders  on  the  variation  of 
crankshaft  torque  is  clearly  shown  in  the  following  table.1 
The  firing  interval  is  constant  in  all  cases. 

TABLE  OF  TORQUE  VARIATION 


Number  of  cylinders.  .  . 

1 

2 

3 

4 

5 

6 

7 

8 

9 

10 

12 

16 

18 

Ratio  of  maximum  in- 

stantaneous   torque 

to  mean  torque  

7.70 

5.20 

2.74 

2.94 

1.64 

1.17 

1.45 

1.40 

1.22 

1.12 

1.13 

1.06 

1.03 

Relative  values  of  the 

maximum  torque  at 

propeller  

1.0 

1.35 

1.07 

1.53 

1.06 

0.91 

1.32 

1.45 

1.43 

1.46 

1.76 

2.20 

2.41 

The  relative  values  of  the  maximum  torque  are  also  given  in 
the  table,  the  value  for  a  single  cylinder  being  taken  as  unity. 
1  From  article  by  G.  D.  ANGLE,  Aviation,  Oct.  1,  1919. 


48 


THE  AIRPLANE  ENGINE 


It  will  be  seen  that  the  maximum  torque  at  the  propeller  end  of  a 
6-cylinder  engine  is  less  than  the  maximum  torque  exerted  by  a 


IWU 
800 
600 

: 

/ 

*•*-*  ^ 

CQ 

4000  ~ 

r 

[s 

3000  '±. 
2000  S. 
1000  t 
0       ^ 
1000  .8 
2000  ^ 
3000^ 
4000  £ 

/  1 

r 

\ 

\ 

f 

/ 

\ 

\i 

\v 

i 

\-jrt 

1 

f 

^ 

i 

v, 

" 

1 

\ 

1 

/ 

\ 

/• 

•^ 

/ 

/ 

\ 

1 

\ 

/ 

/ 

-400 

V 

J 

/ 

^. 

vv 

N 

^ 

1 

x 

f 

' 

s 

x 

/ 

1: 

300 

10°°0                    18 

}                   360                    540                  720 

Crank  Angle, Degrees 
FIG.  36. — Side  thrust  against  the  cylinder  walls  of  the  Liberty-12  engine. 

single  cylinder  and  consequently  the  crankshaft  must  be  as 

strong  at  the  free  end  as  at  the  propeller  end. 

Side  Thrust. — Side  thrust  of  the  piston 
against  the  cylinder  wall  exists  in  con- 
sequence of  the  obliquity  of  the  con- 
necting rod;  it  disappears  at  the  two 
dead  centers.  Its  magnitude,  G  (Fig.  32), 
is  given  by  G  =  P  tan  6.  Since  r  sin  a  = 
I  sin  6, 

G  _        r  sin  a 

P  ~  \/Z2-r2sm2a 

For  the  Liberty  engine  with  indicator 
diagram,  as  in  Fig.  30,  the  side  thrust  is 
as  shown  in  Fig.  36,  the  maximum  value 
reaching  nearly  1,000  Ib.  Change  of  sign 
indicates  change  of  thrust  from  one  side 
of  the  cylinder  to  the  other.  In  sta- 
tionary cylinder  engines  side  thrust  is 
important  only  in  relation  to  frictional 
wear. 

Offset  Cylinders.-The  side  thrust 
during  the  expansion  stroke,  and  the 

friction  loss  and  ^ter  wear,  may  be 
reduced  by  offsetting  the  cylinder:  that  is, 
by  so  locating  it  that  its  center  line  does  not  pass  through  the  axis 
of  the  crankshaft.  Minor  effects  of  this  arrangement  are  that 


FIG.  37. — Diagram  show- 
ing the  effects  of  the  obli- 


ENGINE  DYNAMICS  49 

the  piston  stroke  is  slightly  greater  than  twice  the  crank  throw 
and  the  mean  speed  of  the  piston  is  greater  during  the  down 
stroke  than  the  up  stroke.  This  last  point  has  the  advantage 
of  reducing  the  heat  loss  to  the  cylinder  walls  during  the 
expansion  period. 

In  Fig.  37,  if  k  =  offset  and  x  =  distance  from  a  point  on  the 
piston  to  a  horizontal  plane  through  the  crankshaft  axis  (in 
a  vertical  engine),  then — 

x      =  r  cos  6    +  I  cos  a 

dx  da 

-^r  —   —  r  sin  0  —  I  sin  a  -=—, 
do  bd 

d2x  d2a      ,  /da\2 

—  =  —  r  cos  o  —  I  sin  a  ~rr^  — *  I  (  ~JT  )  cos  a 

db  db  \do/ 

.  /r  sin  b  —  k\ 
a      =  sin-1 1  —  — -j —   — 1 

da        r  cos  b 
db        I  sin  a 

—  I  cos  a-r  sin  b  +  r  cos  b-l  sin  a  -j 
d2a  _ dp 

db2  =  I2  cos2a 

dzx  ,       ,  r2cos26 

— -  —  —  r  cos  b  —  I  cos  a  ^ ^~ 

db2  I2  cos2a 

,  ,   .       da 
I  cos  a-r  sin  b  +  r  cos  b-l  sin  a--^- 


sn  a 


cos    a 


r2  cos2  6    ,  r2  cos26-tan2a 

=  —  r  cos  o  --  j  —       —  r  T  sin  o-  tan  a  -- 

I  cos  a  L  cos  a 

,       r2  cos2  b  ,.,  „  x  .    ,  , 

=  —  r  cos  6  —  T—      -  (1  +  tan2a)  +  r  sm  D-tan  a. 
Z  cos  a  v 

The  acceleration  of  the  reciprocating  parts  is  equal  to 
/27rn\  2  d*x 

(w)  *» 

and  the  accelerating  force  is 


The  side  thrust  is  given  by  G  =  P  tan  a. 

In  using  the  above  equations  it  should  be  noted  that  the  angle 
a  is  positive  when  the  connecting  rod  swings  away  from  the 
crankshaft,  as  in  the  position  shown  in  Fig.  37,  and  becomes 


50 


THE  AIRPLANE  ENGINE 


negative  on  the  other  side  of  the  vertical.     The  angle  a  is  found 
for  any  value  of  b  from  the  equation 

I  sin  a  =  r  sin  b  —  k 


20    40    60    80    100  120  140  160   180  200  220  240  260  280  300  320  340  560 

Degrees  Rotation  of  Crank  from  Top  Dead  Center 

FIG.  38. — Effects  of  different  degrees  of  offset  on  the  side  thrust  in  a  single 
cylinder  of  the  dimensions  of  the  Liberty-12  engine. 

The  results  of  an  analysis  of  the  Liberty  engine  with  offsets  of 
0.5,  1.0  and  1.75  in.  gives  side  thrusts  as  shown  in  Fig.  38.  It 
will  be  seen  that  with  an  offset  of  half  the  crank  throw  (1.75  in.) 
the  side  thrusts  during  the  exhaust  and  compression  strokes  are 

nearly  as  high  as  the  maximum 
value  reached  in  an  engine  with- 
out offset.  The  lowest  maximum 
is  obtained  with  an  offset  of  1  in. 
Rotary  Engines. — The  turning 
moment  in  a  rotary  engine  results 
entirely  from  side  thrust  on  the 
cylinder  walls.  This  thrust  is  due 
not  only  to  the  obliquity  of  the 
connecting  rod,  as  with  stationary- 

FIG.  39. — Diagram  of  rotary  engine.         T    j  •  i      A     i  J.T, 

cylinder  engines,  but  also  to  thrust 

resulting  from  tangential  acceleration  of  the  reciprocating  parts. 
The  radial  and  tangential  accelerations  of  these  parts  and  the 
inertia  forces  resulting  from  them  may  be  determined  as  follows : 

Take  a  single  cylinder,  as  in  Fig.  39,  rotating  about  the  shaft 
0,  while  the  connecting  rod  rotates  about  the  fixed  crankpin  P. 


ENGINE  DYNAMICS  51 

The  piston  pin  Q  will  move  along  the  axis  OX  as  that  axis 
rotates  about  Q.  If  the  length  OQ  =  x,  the  point  Q  will  undergo 
radial  acceleration  aR  along  OQ  and  also  tangential  acceleration 
aT  at  right  angles  to  OQ.  The  magnitudes  of  these  accelerations 
are  given  by  the  general  theorems  : 

da  2 


and 

_     dx  da         d*a 

aT  "  **&<si!+*ifr 

With  rotary  engines  the  angular  velocity  co  of  the  cylinders  is 
constant;  or 

da 

dt"="- 
and 

^  =  0 
dt2 

Consequently 

d2x 

a«  --  W 
and 

.    dx 


To  find  values  of  -j.  and  -^  ,  the  relations,  r  sin  a  =  I  sin  6, 

and,       x  =  r  cos  a  +  I  cos  b,  are  used.     Combining  these, 

r2sin2  a 


x  =  r  cos  a 


r  cos  a  +  I  (1  -      — ^2 — )  approximately.     Then 

\  £jl     I 


x                  .    I       r 
=  cos  a  H =y 


But  a  =  cot 
therefore 


x       I   .  r    .   9 

=  -  +  cos  w^  —  ^  sm2  wt 

T          T  ZL 

1  dx  r    . 

37  =  —  a>  sm  ut  —  co  ,  sm  cof  cos  co^ 
r  dt  I 

1  d2#  T 

^r^  =  —  w2  cos  co^  +  co2,  (sin2wf  —  cos 


52 


THE  AIRPLANE  ENGINE 


Then, 


OR  _  I  d?x       x 


=  --  o>2  —  2co2  cos  ut  +  »co2y  sin2  otf  —  co2,  •  cos2  cof 

7*  2      I  I 


and 


dx 


=  —  2co2r  sin  co£  (  1 


=-  cos  con 


The  values  of  aR,  multiplied  by  the  mass  of  the  reciprocating 
parts,  give  the  forces  necessary  to  overcome  inertia  in  the  radial 
direction.  Combining  these  forces  with  the  total  gas  pressures 
(as  in  Fig.  31)  gives  the  axial  force  P  from  which  the  side  thrust 
G  (Fig.  32)  resulting  from  obliquity  of  the  connecting  rod  is 
obtained,  as  on  page  48.  To  find  the  total  side  thrust  there 
must  be  added  to  this  the  thrust  due  to  the  tangential  accelera- 
tion aTy  which  is  the  product  of  the  acceleration  by  the  mass 
of  the  reciprocating  parts.  The  product  of  the  total  side  thrust 
by  the  distance  of  the  piston  pin  from  the  crankshaft  is  the 
turning  moment. 

For  several  cylinders  the  turning  moments  are  additive.  As 
the  cylinders  are  spaced  at  equal  angular  intervals,  the  total 
turning  moment  at  any  angular  displacement  a  of  one  cylinder  is 


De3ree&  The  quantities  in*  brackets 

are  the  angular  displacements 
of  the  various  cylinders:  n  is 
the  total  number  of  cylinders. 
Mayer1  has  calculated  the 
FIG.  40.— Variation  of  radial  and  tan-  turning    moments    for    vari- 

gential  accelerations  of  the  reciprocating    QUS  arrangements  of  rotating- 
parts  of  a  rotary  engine.  T    j  •  T-I 

cylinder  engines.  For  a  seven- 
cylinder  engine,  110  mm.  diameter,  120  mm.  stroke,  length  of 
connecting  rod  213  mm.,  weight  of  reciprocating  parts  1.3  kg., 
making  1,600  r.p.m.,  the  radial  and  tangential  accelerations  aR 
and  CLT  vary  throughout  the  revolution  as  indicated  in  Fig.  40. 
It  will  be  seen  that  the  radial  acceleration  is  always  negative, 
1(1  Etude  dynamique  des  moteurs  a  cylindres  rotatifs." 


ENGINE  DYNAMICS 


53 


instead  of  changing  sign  as  in  the  acceleration  of  the  recipro- 
cating parts  of  stationary-cylinder  engines.  Combining  the 
inertia  force  required  to  give  the  reciprocating  parts  this  radial 
acceleration  with  the  gas  pressure  force,  the  resultant  force 
acting  along  the  cylinder  axis  varies  as  shown  in  the  solid  black 
line  in  Fig.  41.  It  will  be 


seen    that    this    force    is         I/ 

positive    only  for   a   very           | 

short  period  at  the  begin-    300°  || 

k 

ning  of  expansion.    If  the 

i 

speed    of    the    engine     is    2000-] 

\l 

increased  the  condition  is 

\ 

soon  reached  when  the  re-    iooo-JN 

\ 

\ 

suit  ant  force  on  the  piston            | 

\. 

Degrees                       / 

pin  is  negative  throughout       Q 

, 

\|80 

360             540   _<,/  1 

the    cycle.     In    that  case 

^ 

^ 

s~ 

*^^ 

]  1   1 

the  connecting  rod  would 

Jl 

LUi 

be   under    tension  all  the  ~1000  |j 

/ 

f' 

;\ 

f\ 

time  (a  very  favorable  con- 

) 

j 

<? 

\ 

\ 

//      \  j 

dition,  permitting   consid-  -wotn 
erable    lightening    of    the          L 

I 

'4 

^ 

\ 

^Sw 

> 

/<?         \v/ 

//      \[ 

>« 

rod)    and   the    connecting  -sooo-J 

rod   might   be  replaced  by    FlQ-  41.— Forces  acting  along  the  cylinder 
,     .  ,        .  axis  of  a  rotary  engine. 

a    chain    except    that    it 

is  under  compression   before  the  engine  has  attained  its  full 

speed. 

The  turning  moment  for  a  seven-cylinder  Gnome  engine  of  the 
dimensions  given  above  is  shown  in  Fig.  42.  The  maximum 
excess  of  power  developed  over  the  mean  resistance  (shaded 


720 


!&0 


540 


360 
Degrees 

FIG.  42. — Turning  moment  of  a  7-cylinder  Gnome  engine. 


area)  is  60  ft.-lb.,  or  about  Ml  20  of  the  total  kinetic  energy  of  the 
engine,  and  would  be  a  much  smaller  fraction  of  the  kinetic 
energy  of  engine  and  propeller  combined.  The  earlier  Gnome 
engines  governed  by  cutting  out  the  explosion  on  one  or  more  of 
the  cylinders  according  to  the  power  requirements.  The  cutting 


54  THE  AIRPLANE  ENGINE 

out  of  a  cylinder  has  a  serious  effect  on  the  uniformity  of  turning 
moment  and  results  in  a  maximum  deficiency  of  work  done 
during  the  cycle  of  280  ft.-lb.  as  compared  with  the  maximum 
excess  (or  deficiency)  of  60  ft.-lb.  when  all  cylinders  are 
functioning. 

The  resultant  force  on  the  crankpin  of  the  above  engine  at 
full  load  varies  from  2,640  to  800  Ib. 

BALANCING 

The  forces  acting  in  an  engine  are  of  two  kinds,  in  so  far  as 
their  effect  on  stresses  is  concerned.  The  gas  pressure  in  the 
cylinder,  being  exerted  both  on  the  cylinder  head  and  on  the 
piston,  subjects  the  engine  to  equal  and  opposite  forces  which 
are  inherently  balanced,  or,  in  other  words,  has  no  effect  in 
displacing  the  engine  as  a  whole.  The  inertia  forces  of  the 
moving  part  in  each  cylinder  are  not  inherently  balanced.  If 
the  engine  is  to  operate  smoothly  (without  vibration)  it  must  be 
so  designed  as  to  make  these  unbalanced  forces  counteract  one 
another  as  far  as  possible. 

There  are  two  kinds  of  moving  parts  to  be  considered:  (a) 
the  rotating  parts,  and  (6)  the  reciprocating  parts. 

Rotating  Parts.*- A  body  of  mass  m  (weight  W  =  mg)  revolving 
with  angular  velocity  co  radians  per  second  (n  r.p.m.)  in  a  circle 
of  radius  r  feet  has  an  acceleration  of  co2r  which  gives  rise  to  a 
centrifugal  force  F  =  mco2r  =  0.00034  Wn2r  Ib.  acting  radially. 
The  rotating  body  is  usually  composed  of  the  crank,  crankpin, 
and  the  large  end  of  the  connecting  rod.  The  resulting  force  F 
is  an  unbalanced  force  acting  on  the  crankshaft.  In  multicrank 
engines  there  will  always  be  a  number  of  these  forces,  one  acting 
at  each  crank.  If  there  is  more  than  one  connecting  rod  attached 
to  the  crank  (as  in  Vee  and  fixed  radial  engines)  the  revolving 
mass  includes  the  large  ends  of  all  the  connecting  rods  attached 
to  the  crank.  In  rotating  cylinder  engines  the  revolving  mass 
includes  the  cylinders  themselves  but  not  the  crank  and 
crankpin. 

It  is  easy  to  balance  the  rotating  parts.  To  accomplish  this 
it  is  necessary  (1)  that  the  vector  sum  of  these  unbalanced  forces 
should  be  zero  and  (2)  that  the  sum  of  their  couples  about  any 
(arbitrarily  selected)  plane  should  be  zero.  As  the  centrifugal 
forces  on  all  the  cranks  are  equal,  the  condition  (1)  is  met,  with 
any  number  of  cylinders  greater  than  one,  by  making  the  crank 


ENGINE  DYNAMICS  55 

angle  intervals  equal.  Condition  (2)  can  be  met,  with  any  num- 
ber of  cylinders  greater  than  two,  if  the  cranks  are  spaced  at 
equal  distances  apart  along  the  shaft. 

There  is  an  unbalanced  centrifugal  pressure  on  the  crankpin 
due  to  the  inertia  of  the  big  ends  of  the  connecting  rods,  and  on 
the  main  bearings  due  to  the  inertia  of  all  the  rotating  parts. 

Reciprocating  Parts.-K-The  inertia  force  of  the  reciprocating 
parts  has  already  been  seen  to  be 

Pa  =  0.00034 Wn*r  (cos  a  ±  ^-cos 

If  the  connecting  rod  were  infinitely  long  the  expression  would  be 
Pal  =  0.00034  Wn2r-cos  a. 

The  second  term  in  the  brackets  is  due  to  the  obliquity  of  the 
connecting  rod  and  increases  in  value  as  I  becomes  shorter. 
It  is  customary  to  separate  the  two  terms  in  a  discussion  of 
balancing,  the  quantity  Pal  being  called  the  primary  inertia  force, 
while 

Pan  =  ±  0.00034 Wn*j- cos  2a 

is  called  the  secondary  inertia  force.  It  is  found  that  the  balance 
of  the  primary  forces  is  much  more  easily  achieved  than  that  of 

the  secondary  forces.     Since  j  is  usually  about  J^,  the  magnitude 

of  the  secondary  forces  is  only  about  one-quarter  that  of  the 
primary  forces  so  that  secondary  balance  is  not  as  important  as 
primary  balance. 

The  conditions  for  balance  of  primary  and  secondary  inertia 
forces  are  exactly  the  same  as  those  for  centrifugal  forces.  There 
is,  however,  this  difference,  that  these  forces  act  always  in  the 
direction  of  the  line  of  stroke.  The  practice  of  balancing  the 
primary  forces  by  the  use  of  counterbalancing  masses  attached  to 
the  crankshaft  is  inadmissible  (where  avoidable)  because  of  the 
necessity  of  keeping  the  total  weight  as  low  as  possible  and  also 
because  the  same  result  can  be  obtained  by  multicylinder 
construction. 

The  following  general  results  of  an  analysis  of  various  possible 
cylinder  arrangements  may  be  stated,  it  being  assumed  that  the 
reciprocating  masses  are  the  same  in  all  cylinders: 

With  cylinders  in  a  line  and  equally  spaced  at  intervals  of 
Lft.: 


56  THE  AIRPLANE  ENGINE 

For  three-crank  engines  with  cranks  at  120  deg.,  the  primary 
and  secondary  forces  are  balanced,  but  the  primary  and  secondary 
couples  are  not.  The  maximum  unbalanced  primary  couple  is 
\/3'mco2rL. 

For  four-crank  engines  with  cranks  at  180  deg.,  the  primary 
forces  are  balanced  but  the  secondary  forces  are  not.  The 
secondary  forces  in  a  vertical  engine  have  a  vertical  resultant 

r2 

which  is  equal  to  4mco2  y-  at  each  quarter  turn  (0  deg.,  90  deg.,  180 

deg.,  etc.)  and  become  zero  at  each  eighth  turn  (45  deg.,  135 
deg.,  etc.). 

For  five-crank  engines  with  cranks  at  72  deg.,  the  primary  and 
secondary  forces  are  balanced,  but  the  couples  are  not.  The 
maximum  unbalanced  primary  couple  is  2.5  mco2rL;  the  maximum 

r2 

unbalanced  secondary  couple  is  5mco2y-  L. 

For  six-crank  engines  with  cranks  at  120  deg.,  the  primary  and 
secondary  forces  and  couples  are  all  balanced. 

With  opposed  cylinder  engines,  with  two  cylinders  and  cranks 
at  180  deg.,  the  primary  and  secondary  forces  are  balanced;  the 
primary  couple  is  unbalanced  with  a  maximum  value  of  ma>VL. 

With  Vee  engines  : 

For  six  cylinders,  angle  of  Vee  120  deg.,  three  cranks  at  120 
deg.,  the  primary  and  secondary  forces  are  balanced;  the  couples 
are  unbalanced;  maximum  unbalanced  primary  couple  is 


For  eight  cylinders,  angle  of  Vee  90  deg.,  four  cranks  at  180°, 
primary  forces  balanced,  secondary  forces  unbalanced;  primary 
and  secondary  couples  balanced.  Unbalanced  secondary  force 
acts  always  at  right  angles  to  the  longitudinal  plane  of  symmetry 

r2 

and  has  a  maximum  value  of  5.6  mco2y  • 

For  twelve  cylinders,  angle  of  Vee  60  deg.,  six  cranks  at  120  deg., 
primary  and  secondary  forces  and  couples  all  balanced. 

With  fixed  radial  engine  of  k  cylinders  with  all  the  connecting 
rods  on  one  crank,  the  inertia  forces  have  a  constant  resultant 

k 
of  ^wcoV,  approximately,  along  the  crank.     It  can  be  balanced 

~L*yY} 

completely  by  a  counterbalance  mass  of  —  „-  at  radius  r,  opposite 


ENGINE  DYNAMICS  57 

the  crankpin.  If  there  are  two  banks  of  cylinders,  each  with  k 
cylinders  and  with  cranks  at  180  deg.,  the  primary  forces  will 
balance  but  there  will  remain  an  unbalanced  primary  couple  of 

<-mco2rL. 

z 

Rotating-cylinder  engines  have  two  sets  of  rotating  masses: 

1.  The  cylinders,  which  are  perfectly  balanced,  and 

2.  The  pistons  and   connecting  rods,  which  have  primary 
balance  but  not  secondary  balance. 

PERIODIC  UNBALANCED  FORCES  V 

The  results  of  the  existence  of  periodic  unbalanced  forces 
in  an  airplane  engine  may  be  two-fold:  (1)  to  move  the  engine  as 
a  whole,  and  (2)  to  distort  the  engine. 

The  engine  in  an  airplane  is  supported  on  wooden  members 
which  are  flexible  and  are  in  no  way  the  equivalent  of  a  rigid 
foundation.  If  the  engine  moves  as  a  whole  it  will  flex  its  supports 
instead  of  moving  the  airplane  as  a  whole.  The  maximum 
possible  duration  of  the  periodic  variation  of  any  unbalanced 
force  is  one  engine  revolution  or  about  ^5  sec.;  ordinarily  it  will 
be  not  more  than  one-quarter  of  a  revolution  or  Hoo  sec.  This 
time  is  too  short  to  permit  the  unbalanced  force  to  produce 
appreciable  deviation  of  the  airplane  as  a  whole  although  it 
may  set  up  vibration  in  it. 

The  crankcase  of  the  airplane  engine  is  always  very  thin 
and  consequently  flexible.  Much  of  the  unbalanced  force  may  be 
taken  up  in  producing  distortion  of  the  crankcase.  This  is 
very  undesirable  since  it  necessarily  results  in  varying  align- 
ments of  the  crankshaft  bearings  and  consequent  increase  in 
shaft  friction.  -  . 

Both  the  engine  supports  and  the  crankcase  have  their  own 
natural  periods  of  vibration.  If  the  frequency  of  the  distur- 
bances due  to  unbalanced  forces  in  the  engine  is  the  same  as  the 
natural  frequency  of  the  vibration  of  either  of  these  members,  the 
amplitude  of  vibration  will  be  greatly  increased  beyond  the  very 
small  amount  due  to  a  single  application  of  the  unbalanced  force. 
It  is  most  important  that  the  critical  speed  at  which  the  vibration 
of  the  engine  on  its  supports  becomes  a  maximum  should  be 
avoided,  as  well  as  simple  multiples  of  those  speeds.  The  natural 
periods  of  vibration  of  such  complex  structures  as  engines,  on 
their  supports  in  an  airplane,  can  not  be  calculated.  If  excessive 


58 


THE  AIRPLANE  ENGINE 


vibration  is  found  at  or  near  the  designed  speed  of  the  engine 
the  structures  must  be  altered  to  change  the  natural  period. 
Various  devices  have  been  used  for  neutralizing  the  unbalanced 
secondary  forces  in  four-  and  eight-cylinder  engines.  A  good 
example  of  these  is  the  Lanchester  balancer  (Fig.  43),  which  has 
had  considerable  application  to  automobile  engines.  This 
consists  of  two  exactly  similar  unbalanced  cylinders  located  under 
the  center  main  bearing  and  driven  by  a  gear  on  the  crankshaft. 
These  cylinders  revolve  in  opposite  directions  and  their  unbalanced 
weights  are  located  so  that  their  common  center  of  gravity  travels 
up  and  down  in  a  vertical  plane  and  so  balances  the  displacement 
of  the  common  center  of  gravity  of  the  pistons,  which  falls  in  a 
plane  at  the  middle  of  the  piston  travel  when  the  pistons  are  on 


Mid-stroke 


FIG.  43. — Lanchester  balancer. 

their  dead  centers,  but  falls  below  it  when  the  cranks  have  revolved 
90  deg.  past  that  position. 

It  should  be  clearly  recognized  that  complete  balance  of 
primary  and  secondary  forces  and  couples  does  not  ensure  absence 
of  vibration.  Such  complete  balance  means  that  the  engine  as  a 
whole  has  no  tendency  to  move,  but  there  are  always  internal 
stresses,  particularly  those  imposed  by  the  opposing  couples  on 
the  engine  structure.  If  the  periodicity  of  the  application  of  these 
stresses  coincides  with  the  natural  period  of  the  structure  (or 
some  fraction  of  it)  severe  vibrations  may  be  set  up.  In  any  case 
heavy  bearing  loads  are  likely  to  be  imposed,  especially  in  the 
center  main  bearing,  through  the  action  of  opposing  couples. 

For  torsional  oscillations  of  the  crankshaft  see  p.  146. 

The  following  table  gives  values  of  inertia  and  centrifugal 
forces,  resulting  bearing  pressures  and  other  calculated  quantities 


ENGINE  DYNAMICS 


59 


for  the  six-cylinder  200-h.p.  Austro-Daimler  engine,  the  six- 
cylinder  270-h.p.  Basse-Serve  engine  and  the  12-cylinder  400-h.p. 
Liberty  engine: 

INERTIA  FORCES,  BEARING  LOADS,  ETC. 


Austro- 
Daimler 
engine 

Basse- 
Selve 
engine 

Liberty- 
12 
engine 

Weight  of  piston  complete  with  rings  and  piston  pin,  Ib. 

4.18 
0  188 

6.187 
0  211 

3.838 
0  1955 

Weight  of  connecting  rod  complete  Ib                ... 

4  84 

9  00 

2  9  and 

Weight  of  reciprocating  part  of  connecting  rod,  Ib  
Total  reciprocating  weight  per  cylinder,  Ib  

1.66 
5  84 

2.25 

8  437 

6.35 
1.225 
5  063 

Weight  per  sq.  in.  of  piston  area,  Ib  
Length  of  connecting  rod  (centers),  in  

0.263 
12.40 

0.288 
14  17 

0.258 
12  0 

Ratio  connecting  rod  length  to  crank  throw  

3.6:1 

3.6:1 

3.44:1 

Inertia,  Ib.  per  sq.  in.  piston  area  (top  center)  

63.8 

80.7 

117.0 

Inertia,  Ib.  per  sq.  in.  piston  area  (bottom  center)  

36.2 
25  0 

45.7 
31  6 

65.2 

3  18 

6  75 

1  675  and 

Total  centrifugal  pressure,  Ib  

610  0 

1  480  0 

5.125 
1,469  0 

Centrifugal  pressure,  Ib.  per  sq.  in.  piston  area  
Mean  average  loading  on  crankpin  bearing,  total  from 
all  sources,  Ib.  per  sq.  in.  piston  area  
Diameter  of  crankpin,  in  
Rubbing  velocity,  ft.  per  sec  

27.5 

91.0 
2.20 
13.42 

50.7 

115.7 
2.75 
16.8 

74.8 

175.0 
2.375 
17.7 

Effective  projected  area  of  big-end  bearing,  sq.  in  
Ratio  piston  area  to  projected  area  of  big-end  bearing  . 
Mean  average  loading  on  big-end  bearing,  Ib.  per  sq.  in. 

5.02 
4.42:1 
402.0 

5.39 
3.5:1 
405.0 

5.34 
3.68:1 
642.0 

CHAPTER  IV 
ENGINE  DIMENSIONS  AND  ARRANGEMENTS 

Certain  special  requirements  control  the  selection  of  the 
dimensions  and  arrangements  of  airplane  engines.  These  are: 

1.  Minimum  total  weight  of  the  engine  and  its  accessories   per  brake 
horse  power  developed,  or  maximum  power  output  per  pound  of  weight. 

2.  Maximum  fuel  economy. 

3.  Compactness. 

4.  Freedom  from  unbalanced  forces  and  from  vibration. 

5.  Reliability. 

To  obtain  minimum  weight  per  horse  power,  it  is  necessary  that 
the  engine  should  have  minimum  weight  per  cubic  foot  of  piston 
displacement  per  revolution,  and  that  it  should  be  operated  with 
maximum  power  per  cubic  foot  of  cylinder  volume.  The  latter 
demands  a  combination  of  the  maximum  obtainable  mean 
effective  pressure  with  high  speed  of  revolution.  The  mean 
effective  pressure  is  constant  at  moderate  engine  speeds  but  falls 
off  at  high  speeds  in  consequence  of  falling  volumetric  efficiency. 
Beyond  a  certain  limiting  speed  the  mean  effective  pressure  will 
fall  off  more  rapidly  than  the  increase  in  engine  revolutions  per 
minute  (see  p.  37)  and  the  engine  power  will  decrease.  This 
limiting  speed  should  be  made  as  high  as  possible  by  making 
the  valve  openings  large  and  the  inlet  and  exhaust  manifolds  short 
and  of  ample  cross-section. 

All  airplane  engines  must  be  multicylindered  in  order  to  give 
the  necessary  uniformity  of  turning  moment  and  freedom  from 
unbalanced  forces.  The  weight  of  the  engine  per  cubic  foot  of 
piston  displacement  per  revolution  will  depend  on  the  unit  size 
of  cylinder  selected.  Comparing  two  unit  cylinders  of  exactly 
the  same  form  but  of  different  sizes,  it  will  be  found  that  the 
thickness  of  the  cylinder  walls  will  not  have  to  be  increased  as 
rapidly  as  the  cylinder  diameter  because  the  wall  (for  structural 
reasons)  is  always  made  thicker  than  the  stresses  demand,  by  an 
amount  which  does  not  vary  much  with  the  diameter.  Conse- 
quently the  weight  of  the  cylinder  per  cubic  foot  of  piston  dis- 

60 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS  61 

placement  diminishes  with  increased  size,  and  the  same  is  true 
of  most  of  the  other  engine  parts. 

On  the  other  hand,  the  weight  of  the  engine  per  cylinder 
diminishes  with  increase  in  the  number  of  cylinders  in  line.  The 
engine  consists  of  a  number  of  exactly  similar  units  (cylinder, 
running  parts,  section  of  crankcase,  etc.)  and  certain  approxi- 
mately constant  weights  such  as  ends  of  crankcase,  pumps, 
magnetos,  propeller  hub  and  so  forth.  The  addition  of  more 
cylinders  will  diminish  the  weight  of  the  engine  per  cylinder. 

A  further  diminution  in  weight  of  the'  engine  per  cylinder 
can  be  obtained  by  using  the  Vee  or  W  arrangement  of  cylinders. 
A  substantial  saving  in  the  weight  of  crankshaft  and  crankcase, 
per  cylinder,  results  from  these  arrangements.  Still  greater 
saving  results  from  the  adoption  of  the  radial  arrangement. 

Considerations  of  torque  and  balance  (see  p.  56)  indicate 
that  the  number  of  cylinders  should  not  be  less  than  six  for 
an  engine  of  moderate  power  (100  to  150  h.p.).  For  higher 
powers  the  choice  is  between  more  cylinders  and  larger  cylinders. 
Engines  have  been  built  with  as  many  as  24  cylinders,  but  it  does 
not  seem  likely  that  this  number  will  be  much  used  or  exceeded. 

Cylinder  size  can  be  increased  by  increasing  either  diameter  or 
stroke  or  both.  The  diameter  is  at  present  limited  to  from  6  to 
7  in.  by  the  difficulty  of  keeping  the  piston  cool.  The  heat 
given  to  the  center  of  the  piston  has  to  travel  radially  to  the 
walls,  and  it  is  necessary  to  increase  the  thickness  of  the  piston  as 
the  diameter  increases  in  order  to  give  sufficient  section  of  metal 
to  carry  the  heat  away.  This  results  in  a  heavy  piston  and 
excessive  inertia  forces  in  the  reciprocating  parts.  The  engine 
stroke  is  also  limited  at  present  to  8  in.;  increase  in  stroke  beyond 
that  limit  increases  the  over-all  height  of  the  engine  to  dimensions 
which  are  difficult  to  accommodate  without  increasing  the  size 
of  the  fuselage.  Furthermore,  increase  in  stroke  increases  the 
weight  of  the  engine  more  than  does  a  corresponding  increase  in 
diameter.  The  small  ratio  of  stroke  to  diameter  which  character- 
izes airplane  engines  is  not  as  objectionable  as  it  would  be  in 
other  engines;  it  increases  the  ratio  of  water-jacketed  surface  to 
cylinder  volume,  but  the  percentage  of  heat  lost  to  the  jacket  is 
nevertheless  smaller  than  in  other  engines  in  consequence  of  the 
high  engine  speed  and  high  mean  effective  pressure. 

The  weight  of  the  engine  per  cubic  foot  of  piston  displacement 
per  revolution  is  also  a  function  of  the  ratio  of  connecting  rod 


62  THE  AIRPLANE  ENGINE 

length  to  stroke.  The  smaller  this  ratio  the  less  is  the  over-all 
height  and  the  weight  of  the  engine.  The  objection  to  a  small 
ratio  is  the  increase  in  magnitude  of  the  secondary  inertia  forces 
which  result  from  the  obliquity  of  the  connecting  rod.  These 
secondary  forces  can  be  perfectly  balanced  with  certain  arrange- 
ments of  cylinders  (see  p.  56)  and  the  objection  eliminated. 
The  ratio  usually  ranges  from  1.5  to  1.7. 

High  fuel  economy  is  important  primarily  in  its  effect  on  the 
weight  to  be  carried  by  the  plane.  For  a  five-hour  flight,  an  engine 
weighing  2.5  Ib.  per  Horse  power  and  using  0.5  Ib.  of  fuel  per  horse- 
power hour  will  have  the  same  total  weight  of  engine  and  fuel 
as  an  engine  weighing  2  Ib.  per  horse  power  and  using  0.6  Ib.  of 
fuel  per  horse-power  hour.  The  heavier  and  more  efficient 
engine  would  ordinarily  be  the  better  of  the  two  in  respect  to 
reliability  and  durability.  To  obtain  high  fuel  economy  the 
most  important  factor  is  the  ratio  of  compression  which  should 
be  as  great  as  can  be  used  without  detonation  or  preignition. 

Compactness  is  important  both  in  respect  to  frontal  area  and 
over-all  length.  The  frontal  area  of  large  engines  should,  if 
possible,  be  of  such  form  and  dimensions  as  not  to  require  any  in- 
crease in  the  cross-section  of  the  containing  fuselage.  Short  over- 
all length  is  distinctly  advantageous  and  permits  the  fuel  tanks 
to  be  located  close  to  the  center  of  gravity  of  the  plane. 

Freedom  from  vibration  is  necessary  because  the  mounting 
of  the  engine  is  on  elastic  supports — usually  wooden  longerons — 
which  leave  the  crankcase  free  to  distort  under  internal  forces. 
Such  arrangements  of  cylinders  as  will  eliminate  or  minimize 
unbalanced  forces  are  desirable  but  they  cannot  be  relied  upon 
to  prevent  vibration.  Even  with  the  most  perfect  balancing, 
torsional  vibration  of  the  crankshaft  may  produce  excessive  vibra- 
tions at  certain  engine  speeds;  such  speeds  must  be  avoided. 

Engine  Arrangements. — Airplane  engines  have  been  built  in  a 
great  variety  of  arrangements  of  which  a  number  have  survived. 
These  may  be  classified  as  (1)  radial,  (2)  vertical,  (3)  Vee,  (4) 
W,  (5)  X.  Radial  engines  (both  fixed  and  rotary)  are  discussed 
in  Chapter  VIII.  They  have  minimum  weight  per  horse  power 
and  shortest  over-all  length,  but  they  have  maximum  frontal 
area  and  until  very  lately  have  shown  economy  inferior  to  that  of 
the  other  types.  They  are  generally  air  cooled. 

The  vertical  engine  (Figs.  8  and  9)  has  one  row  of  cylinders. 
The  Vee  engine  (Figs.  46  and  47)  has  two  linear  rows  of  cylinders; 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


63 


the  "  angle  of  the  Vee"  is  the  acute  angle  between  the  axial  planes 
of  the  two  rows.  The  W  or  \J/  engine  (Figs.  72  and  73)  has 
three  rows  of  cylinders  of  which  the  central  one  is  vertical  and  the 
other  two  form  equal  angles  with  the  vertical.  The  X  engine  has 
four  rows  of  cylinders  arranged  symmetrically  about  the  vertical 
and  horizontal  planes  but  not  necessarily  with  equal  angles 
between  the  planes  of  the  cylinder  axes.  In  the  radial  engine 
(Fig.  136)  three  or  more  cylinders,  constituting  a  group,  have  their 
axes  intersecting  at  a  point  on  a  common  shaft.  The  radial  or 
fixed  engine  should  not  be  confused  with  the  rotary  engine,  in 
which  the  cylinders  revolve  about  a  stationary  crankshaft. 
A  radial  engine  may  consist  of  more  than  one  group  or  bank  of 
cylinders,  each  group  having  its  own  common  plane  of  cylinder 
axes,  perpendicular  to  the  axis  of  the  shaft. 

Each  cylinder  of  a  four-cycle  engine  requires  two  revolutions 
or  720  deg.  of  rotation  of  the  crankshaft  to  complete  its  cycle. 


1234 


/     2     3     4     5      6 


FIG.  44. 
Four-cylinder  engine 


FIG.  45. 
Six-cylinder  engine 


The  explosions  in  a  cylinder  occur  every  other  revolution.  If 
explosions  are  to  occur  in  n  cylinders  at  equal  intervals,  the 
interval  expressed  in  degrees  of  crankshaft  rotation  must  be 
720  -T-  n.  Explosion  always  occurs  when  the  piston  is  closest  to 
the  cylinder  head.  With  a  four-cylinder  engine,  the  interval 
between  explosions  is  180  deg.;  the  crankpins  lie  all  in  one  plane 
passing  through  the  shaft  axis,  one  possible  arrangement  being 
shown  in  Fig.  44.  A  six-cylinder  vertical  engine  has  the  cranks 
720  -f-  6  =  120  deg.  apart:  the  crankpins  lie  in  three  planes 
intersecting  at  the  shaft  axis.  One  arrangement  is  shown  by 
Fig.  45.  Constancy  of  interval  between  impulses  may  be 
obtained  with  crank  dispositions  other  than  those  shown  in 
Figs.  44  and  45.  For  example,  in  Fig.  44,  cranks  2  and  3  may 
be  either  in  line  or  opposed.  If  the  former,  cranks  1  and  4  will 
be. in  line  and  180  deg.  away  from  2  and  3.  If  the  latter,  1  and  4 
will  be  opposed,  and  either  may  be  in  line  with  2.  If  equal  inter- 


64  THE  AIRPLANE  ENGINE 

vals  of  time  between  explosions  were  the  sole  requisite,  the 
number  of  possible  crank  arrangements  with  a  6-cylinder  engine 
would  be  large.  The  actual  crank  arrangements  are  determined 
mainly  by  considerations  of  engine  balance. 

In  a  Vee  engine  each  pair  of  cylinders  (in  one  plane)  acts  on  a 

common  crankpin.     If  there  are  n  cylinders  in  <->  pairs,  the  crank 

interval  is  720  -f-  ~.     The  interval  between  explosions  of  the 

cylinders  is  720  -5-  n.  If  6  is  the  angle  of  the  Vee,  any  one  crank 
moves  through  the  angle  6  in  the  interval  necessary  for  the  two 
pistons  actuating  it  to  reach  respectively  their  highest  positions. 
As  the  explosions  occur  with  the  pistons  in  their  highest  positions, 
the  explosion  in  a  leading  cylinder  will  precede  that  of  its  fol- 
lowing cylinder  by  an  angle  of  0,  or  360  +  6,  deg.  according  to 
whether  both  cylinders  explode  during  the  same  revolution  of  the 
crank  or  the  explosions  occur  in  succeeding  revolutions.  The 
latter  is  always  the  case,  because,  if  the  second  explosion  occurred 
after  the  crank  angle  6,  the  explosion  pressure  would  be  trans- 
mitted to  a  crankpin  which  was  already  subjected  to  the  pressure 
of  the  gas  expanding  in  the  other  cylinder  of  the  pair  and  the 
crankpin  would  be  unduly  loaded.  For  equal  explosion  intervals 
the  angle  0  is  equal  to  720  -f-  n.  This  leads  to  a  90-deg.  angle 
for  8-cylinder  and  a  60-deg.  angle  for  12-cylinder  Vee  engines. 
These  angles  give  the  maximum  uniformity  of  turning  movement; 
other  angles  may  be  employed  for  special  reasons  but  always  at 
some  sacrifice  of  uniformity  of  turning  movement. 

The  choice  as  between  vertical,  Vee  and  W  arrangement  is 
largely  determined  by  the  number  of  cylinders.  It  has  been 
found  that  much  trouble  is  experienced  with  a  crankshaft  having 
more  than  six  cranks  and  the  over-all  length  of  the  engine  becomes 
excessive.  Eight-crank  engines  have  been  built  but  they  have 
not  survived.  The  perfect  balancing  of  the  six-crank  engine 
has  made  it  a  favorite.  With  six  cylinders  the  vertical  arrange- 
ment is  almost  universally  adopted;  it  has  the  advantage  of 
having  minimum  frontal  area  and  consequently  of  being  most 
easily  accommodated  in  the  fuselage.  With  eight  cylinders  the 
90-deg.  Vee  engine  with  four  cranks  is  generally  accepted  as  the 
best  arrangement  although  the  wide  Vee  angle  results  in  con- 
siderable over-all  engine  width.  With  12  cylinders  the  usual 
arrangement  is  a  60-deg.  Vee  and  six  cranks.  The  45-deg.  Vee 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS  65 

adopted  in  the  Liberty  engine  results  in  decreased  width  of 
engine  but  increased  height;  it  also  results  in  a  less  uniform  turn- 
ing moment  on  the  crankshaft. 

The  Warrangement  for  12  cylinders  is  with  three  rows  of  cylinders 
and  four  cranks;  this  shortens  the  over-all  length  and  decreases 
the  weight  but  greatly  increases  the  engine  width,  especially 
if  a  60-deg.  angle  (which  gives  most  uniform  torque)  is  used 
between  the  rows.  With  18  cylinders  the  W  arrangement  with 
three  rows  of  cylinders  and  six  cranks  is  used;  the  angle 
between  the  rows  for  most  uniform  torque  should  be  40  deg., 
which  diminishes  the  over-all  width  as  compared  with  the  12- 
cylinder  W  engine. 

Vertical,  Vee  and  W  engines  are  nearly  always  water- 
cooled  in  airplane  practice.  Air  cooling  has  been  successful  only 
with  low  compression  ratios. 

Engine  Dimensions. — Table  3  gives  the  general  dimensions 
and  arrangement  of  the  principal  American  and  foreign  airplane 
engines.  The  horse  powers  given  are  generally  the  maker's 
rating,  but  they  are  naturally  variable  with  the  fuel,  the  car- 
buretor, the  manifolding,  the  ratio  of  compression,  and  the  engine 
speed.  The  ratio  of  compression  is  readily  varied  by  changing 
the  dimensions  of  the  piston  and  for  certain  engines  different 
pistons  are  supplied,  according  to  whether  the  engine  is  to  be 
flown  at  low  altitude  (as  in  seaplanes)  or  at  higher  altitudes. 
The  dry  weight  includes  carburetors,  magnetos,  and  propeller 
hub.  The  weight  per  horse  power  naturally  varies  with  the 
horse  power  and  is  given  for  the  rated  horse  power.  It  is  the 
product  of  two  factors,  the  weight  per  cubic  inch  of  piston 
displacement  and  the  piston  displacement  per  horse  power. 
The  piston  displacement  per  horse  power  is  the  aggregate  dis- 
placement volume  of  the  cylinders  per  stroke  divided  by  the 
horse  power  developed.  It  is  an  excellent  measure  of  the  degree 
to  which  the  piston  displacement  is  utilized.  The  aggregate 
displacement  volume  is  equal  to  I  X  a  X  n  cu.  in.  where  I  is  the 
stroke  in  inches,  a  the  piston  area  in  square  inches,  and  n  is  the 
,  r  ,  _,  p  X  iXaXnXN 

number  of  cylinders.     The  horse  power  =     2  X  12  X  33  000    * 

where  p  is  the  mean  effective  pressure,  and  N  is  the  revolutions 
per  minute.  The  piston  displacement  per  horse  power  =  33,000 
X  24  -T-  p  X  N;  that  is,  it  depends  only  on  the  mean  effective 
pressure  and  the  revolutions  per  minute. 


66 


THE  AIRPLANE  ENGINE 


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ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


69 


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70 


THE  AIRPLANE  ENGINE 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


71 


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72 


THE  AIRPLANE  ENGINE 

TABLE  4. — DETAILED  DIMENSIONS  OF 


Engine 

Liberty 

Packard 

Approximate  horse  power 

200 

400 

180 

270 

5.00 
7.00 
6 
1.40 
137.4 
824.5 
31.1 
168.5 
5.42 
18.46 
7 
(  6-0.  998 
\  1-1.1235 

1 
2.50 
He 
30° 
3.202 
0.013-0.016 

1 
2.50 
0.375 
30° 
2.748 
0.019-0.021 

2 

62.2 
50.0 

85.7 
71.5 

26.5 
32.9 
31.8 

10° 

45° 

48° 
8° 

Ribless 

19.63 
Aluminum 

3.469 
5.469 
3.75 

0.0395 
0.020 

5.00 
7.00 
12 
1.40 
137.4 
1650.0 
31.1 
168.5 
5.42 
18.46 
7 
6-0.998 
1-1  .  1235 

1 
2.50 
He 
30° 
3.202 
0.013-0.016 

1 
2.50 
0.375 
30° 
2.748 
0.019-0.021 

2 

62.2 
50.0 

85.7 
71.5 

26.5 
32.9 
31.8 

10° 
45° 

48° 
8° 

Ribless 

19.63 
Aluminum 

3.469 
5.469 
3.75 

0.0395 
0.020 

4.75 
5.25 
8 
1.105 
93.0 
744.0 
23.17 
116.2 
5.02 
24.9 
5 
4-0.998 
1-1.124 

1 
2.00 
He 
30° 
2.61 
0.013-0.016 

1 
2.00 
0.375 
30° 
2.20 
0.019-0.021 

2 

63.5 
47.5 

81.5 
61.5 

24.0 
35.0 
35.0 

10° 
45° 

48° 
8° 

Ribless 

17.72 
Al.  alloy 

2.25 
4.25 
3.094 

0.037 
0.017 

4.75 
5.25 
12 
1.105 
93.0 
1116.4 
20.4 
113.4 
5.56 
18.0 
7 
6-0.998 
1-1  .  124 

1 
2.00 
He 
30° 
2.61 
0.013-0.016 

1 
2.00 
0.375 
30° 
2.20 
0.019-0.021 

2 

63.5 
47.5 

81.5 
61.5 

24.0 
35.0 
35.0 

10° 
45° 

48° 
8° 

Ribless 

17.72 
Al.  alloy 

2.67 
4.665 
2.99 

0.034 
0.017 

Stroke-bore  ratio                     

Piston  displacement  per  cylinder,  cu.  in. 
Total  piston  displacement,  cu.  in  

Compression  volume,  cu.  in  
Total  volume  of  cylinder,  cu.  in  
Compression  ratio               

Per  cent  compression     

Camshaft  bearings,  number  on  shaft  .  .  . 

Inlet  valves: 
Number  per  cylinder  
Port  diameter,  inches  

Lift,  inches  
Angle  of  seat     

Total  area  of  opening,  square  inches. 
Tappet  clearance,  inches  
Exhaust  valves: 
Number  per  cylinder  
Port  diameter,  inches  
Lift  inches      

Angle  of  seat          

Total  area  of  opening,  square  inches  . 
Tappet  clearance,  inches  

Valve  springs: 
Number  per  valve     

Tension  inlet  (both  springs): 
Valve  open   pounds  

Valve  closed  pounds     

Tension  exhaust  (both  springs): 
Valve  open,  pounds  

Valve  closed  pounds  .           

Internal  spring  tension: 
Valve  closed  pounds       

Valve  open  (inlet),  pounds  
Valve  open  (exhaust),  pounds  
Valve  timing: 
Inlet: 
Opens,  past  top  center  
Closes   past  bottom  center     

Exhaust: 
Opens  before  bottom  center  

Clones   past  top  center  

Piston: 
Tvoe 

Area  of  head,  square  inches  
Material  
Distance  from  center  of  pin  to  top  of 
piston,  inches    

Length  over  all,  inches  

Length  of  bearing  in  cylinder  

Clearance  in  cylinder: 

Bottom,  inches.  .  . 

ENGINE  DIMENSIONS  AND  ARRANGEMENTS 

MERICAN    AND    GERMAN   ENGINES 


73 


Hispano-Suiza 

Benz 

Maybach 

Austro- 
Daimler 

Basse- 
Selve 

Mercedes 

180 

300 

200 

300 

200 

270 

180 

4.72 

5.511 

5.512 

6.50 

5.31 

6.10 

5.51 

5.11 

5.905 

7.480 

7.09 

6.89 

7.87 

6.30 

8 

8 

6 

6 

6 

6 

6 

1.08 

1.07 

1.357 

1.09 

1.296 

1.29 

1.14 

'    89.9 

140.8 

178.4 

235.3 

152.8 

230.2 

150.3 

719.0 

1126.0 

1070.4 

1410.0 

916.8 

1381.0 

901.7 

20.7 

32.6 

37.2 

47.6 

38.0 

69.0 

41.3 

110.6 

173.4 

215.6 

282.9 

190.8 

299.2 

191.6 

5.33 

5.32 

5.8 

5.95 

5.02 

4.34 

4.64 

18.76 

18.8 

17.25 

16.8 

19.9 

23.0 

21.5 

3 

3 

4 

5 

4 

1.342 

1.344 

0.984 

1 

1 

2 

2 

2 

2 

1 

1.968 

2.205 

1.693 

1.89 

1.73 

2.20 

2.677 

0.393 

0.511 

0.433 

0.372 

0.390 

0.390 

0.453 

45° 

45° 

30° 

30° 

45° 

45° 

1.894 

2.79 

2.214 

2.33 

2.12 

2.72 

3.81 

0.078 

0.030 

0.016 

0.012 

0.01 



0.017 

1 

1 

2 

2 

2 

2 

7 

1.968 

2.205 

1.693 

1.89 

1.73 

2.20 

2.677 

0.393 

0.511 

0.433 

0.368 

0.40 

0.390 

0.453 

45° 

45° 

30° 

30° 

45° 

45° 

1.894 

2.79 

2.214 

2.33 



2.72 

3.81 

0.078 

0.030 

0.016 

0.016 

0.012 

0.014 

2 

2 

1 

1 

1 

1 

1 

75.0 

80.0 

43.0 

133.0 

44.5 

42.0 

25.5 

101.8 

. 

75.0 

80.0 

44.0 

133.0 

44.5 

42.0 

24.0 

101.0 

19.0 

30.0 

30.0 

10° 

—  10° 

5° 

-8° 

-10° 

0° 

50° 

62° 

45° 

35° 

30° 

40° 

45° 

62° 

55° 

33° 

45° 

40° 

10° 

29° 

18° 

7° 

7° 

10° 

Ribless 

Ribless 

Ribbed 

Ribless  flat- 

Ribbed 

Convex 

Steel  concave 

top 

crown 

crown 

17.53 

23.82 

23.82 

33.15 

22.2 

29.2 

23.84 

Aluminum 

Aluminum 

Al.  alloy 

Cast  iron 

Aluminum 

Aluminum 

Cast  iron 

1.765 

3.00 

2.480 

3.19 

2.09 

4.375 

5.12 

4.843 

5.944 

4.35 

5.11 

3.071 

0.040' 

0.040 

0.025 

0.029 

0.039 

0.020 

0.0165 

0.020                 0.008 

0.009 

0.016 

0.01 

74 


THE  AIRPLANE  ENGINE 


Engine 

Liberty 

Packard 

Approximate  horse  power 

200 

400 

180 

270 

Rings  : 
Number  per  piston: 
TOD 

3 
0 
11-15 
0.248 

0.031 

4.234 
1.25 

12.0 

4 
0.311 
^6-24 

3 
0 
11-15 
0.248 

0.031 

4.234 
1.25 

12.0 
2 
0.4365 
K6-20 

12.0 
4 
0.3115 

He-24 
1.714 

2.00 
1.25 
Bronze 

2.25 
2.375 
Bronze 
babbitt 
lined 

3 

0 
9.0 
0.217 

0.031 

4.141 
1.25 

9.0 
2\ 
0.437 
Kc-20 

9.0 
4 
0.312 

Hs-24 
1.714 

2.00 
1.25 
Bronze 

2.492 
2.125 
Bronze 
babbitt 
lined 

1.125 
2.631 
Steel 

5 
1-1  .  625 
2-1.750 
1-2.625 
1-4.438 
2.375 

Packard 

Double 
Venturi 

2.0 
1.34 

0.0730 
0.0700 
0.0785 

3 

0 
9.0 
0.217 

0.031 

4.138 
1.25 

9.0 
2 
0.437 
Kc-20 

9.0 
4 
0.312 

He-24 
1.714 

2.00 
1.25 
Bronze 

2.492 
2.125 
Bronze 
babbitt 
lined 

1.125 
2.631 
Steel 

7 
1-1.625 
4-1.750 
1-3.000 
1-4.438 
2.375 

Packard 

Double 
Venturi 

2.0 
1.42 

0.1285 
0.0995 
0.1040 

Bottom 

Tension  pounds 

Width  inches 

Width  of  gap   (ring  in  cylinder), 
inches 

Pin: 

Length  inches 

Diameter  inches 

Connecting  rods,  plain: 
Length  c  to  c 

Number  of  bolts 

Diameter  of  bolts  inches 

Thread 

Connecting  rods,  forked: 
Length  c  to  c 

Number  of  bolts 

Diameter  inches 

Thread  ... 

Rod-stroke  ratio 

1.714 

2.00 
1.25 
Bronze 
babbitt 
lined 

Wristpin  bearing: 
Length,  inches 

Diameter,  inches 

Material 

Big  end  bearing,  forked: 
Length,  inches 

Diameter,  inches 

Material 

Big  end  bearing,  plain: 
Length,  inches 

2.25 
2.375 
Bronze 
shell  babbitt 
lined 

7 
6-1.625 
1-4.25 

2.625 

Zenith 
55  AS 
Single  nozzle 

2.165 
1.575 

1.80mm. 
1  .  55  mm. 
1.00  mm. 

Diameter,  inches 

Material 

Crankshaft  bearings: 
Number 

7 
6-1  .  625 
1-4.25 

2.625 

Zenith 
U.S.,  No.  52 
Double 
annular 
nozzles 
2.047 
1.417 

1  .  65  mm. 
1.70  mm. 
1  .  00  mm. 

Length,  inches 

Diameter   inches 

Carburetor: 
Name  .... 

Model  .  ,    

Type  ".        

Diameter  of  flange,  inches  

Choke,  inches  

Jets,  diameter: 

Compensating  inches      ...    . 

Idling  well  or  pilot  iet.  inches.  .  . 

ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


75 


Hispano-Suiza 

Benz 

Maybach 

Austro- 
Daimler 

Basse- 
Selve 

Mercedes 

180 

300 

200 

300 

200 

270 

180 

3  and  one 

3 

3 

3 

scraper 

3 

3 

3 

1 

1 

1 

0 

0 

1 

I 

4.0 

14.0 

7.0 

0.0984 

0.25 

0.118 

0.255 

0.275 

0.230 

0.011 

0.018 

0.072 

0.055 

0.019 

4.125 

4.781 

4.764 

6.26 

5.10 

1.181 

1.375 

1.299 

1.496 

1.100 

1.37 

8.858 

10.375 

12.992 

12.205 

12.40 

14.17 

2 

2 

4 

4 

4 

4 

8.0  mm. 

0.500 

0.394 

0.551 

0.39 

0.47 

Pitch, 

Pitch, 

Pitch, 

Pitch, 

1  .  52  mm. 

0.9  mm. 

1  .  5  mm. 

1.75  mm. 

8.858 

10.375 

4 

4 

12  mm. 

0.375 

Pitch, 

1.25  mm. 

1.73 

1.76 

1.736 

1.72 

1.80 

1.80 

2  .  203                  2  .  250 

2.795 

3.66 

2.64 

3.14 

1.183 

1.375 

1.299 

1.496 

1.10 

1.37 

Bronze 

Bronze 

Bronze 

Cast  iron 

Phosphor- 

Phosphor- 

bush 

bronze 

bronze 

2.508 

2.50 

1.97 

2.125 

Bronze 

Bronze 

babbitt 

babbitt 

lined 

lined 

1.25 

1.125 

2.441 

2.893 

2.63 

3.03 

2.50 

2.75 

2.362 

2.598 

2.20 

2.75 

Steel  on 

Steel  on 

Bronze 

Bronze, 

Bronze, 

Bronze, 

bronze 

bronze 

babbitt  lined 

white  metal 

white  metal 

white  metal 

, 

5 

7 

7 

7 

7 

7 

1-3.86 

1-4.34 

1-1.772 

2.638 

1-2.20 

1-2.36 

3-1  .  578 

3-2  .  00 

5-1.693 

1-1.71 

5-2.48 

1-Steel  ball 

1-Steel  ball 

1-2.402 

5-1.97 

1-3.74 

2.282 

2.50 

(1)  -1.890 

2.598 

2.28 

2.75 

(2-7)-2.441 

Stromberg 

Stromberg 

Benz 

Maybach 

Austro- 

Bass6- 

Mercedes 

NAD  4 

NAD  6 

BZ  3A-137 

Daimler 

,.Selve 

Double 

Double 

Barrel 

Venturi 

Venturi 

throttle 

Dual 

Twin  jet  dual 

2.18 

2.375 

1.89 

1.50 

1.812 

1.654 

0.945 

1.96 

0.945 

0.0935 

0.039 

Variable 

0.0102 

1.473  mm. 

0.1286 

0.116 

0.020 

0.0046 

0.558  mm. 

76 


THE  AIRPLANE  ENGINE 


TABLE  4.— 


Engine 

Liberty 

Packard 

Approximate  horse  power 

200 

400 

180 

270 

Ignition: 
TVDC 

Battery 
Delco 

30° 
10° 

1.0  XCS 
0.5XCS 
1-5-3-6-2-4 

Counter 
clock 

2 
Cyl.  head 

18.0 
1.5 
0.015-0.018 
1  .  5  X  CS 
1.5  X  CS 

4.7 
1.1 

Battery 
Delco 

30° 
10° 

1.5  XCS 
0.5XCS 
1L-6R—  5L- 
2R-3L-4R- 
6L-1R-2L- 
5R-4L-3R 
Counter 
clock 

2 

Cyl.  head 

18.0 
1.5 
0.015-0.018 
1  .  5  X  CS 
1.5  X  CS 

4.9 
1.3 

4.4 
1.9 

5.7 
3.7 

8.9 

0.8 

510 
10,000 
4,000 
1,872 
(1-5)1,173 

(6)1,393 
(7)364 

Battery 
Delco 

45° 
10° 

2.0XCS 
0.5XCS 
1L-4R-3L- 
2R-4L-1R- 
2L-3R 

Counter 
clock 

2 
Side  of  head 

18.0 
1.5 
0.015 
1.5  X  CS 
1.5  XCS 

3.9 
1.1 

3.3 
1.4 

4.4 
2.5 
6.9 

0.93 

465 
8,240 
3,290 
1,548 
(1,  2,  4)1,066 

(3)694 
(5)402 

Battery 
Delco 

45° 
10° 

2.0XCS 
0.5XCS 
1L-6R-5L- 
2R-3L-4R- 
6L-1R-2L- 
5R-4L-3R 
Counter 
clock 

2 
Side  of  head 

18.0 
1.5 
0.015 
1.5  X  CS 
1.5  X  CS 

9.2 
1.1 

3.3 

1.4 

4.4 
2.5 
6.9 

0.93 

527 
9,320 
3,730 
1,751 
(1,  2,  3,  5,  6) 
1,207 
(4)682 
(7)455 

Manufactured  by  

Maximum  spark  advance,  before  top 
center  

Maximum  retard,  after  top  center.  .  . 
Speed   ignition    drives  (CS  =  crank- 
shaft speed): 
Generator  or  magneto  

Rotation,   direction   of   (facing  pro- 
peller) 
Spark  plugs: 

Location 

Size,  millimeters 

Pitch   millimeters 

Reciprocating      and      centrifugal 
weights  : 
Piston,    complete   with   rings   and 

Upper  end  connecting  rod,  pounds. 
Lower  end  forked  connecting  rod, 

Lower    end   plain   connecting   rod, 
pounds  
Total   weight    of    connecting  rods 
complete,  pounds: 
Forked  

3.7 

0.0 

4.8 
4.8 

0.8 

510 
10,000 
4,000 
1,872 
(1-5)1,173 

(6)  1,393 
(7)364 

Plain  

Total  

Loads      from      maximum      explosion 
pressure,    pounds   per   square 
inch: 
Assumed     maximum     explosion 

Total  load  on  piston  head,  pounds 

Load  on  main  bearings  

ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


77 


Hispano-Suiza 

Benz 

Maybach 

Austro- 
Daimler 

Basse- 
Selve 

Mercedes 

180 

300 

200 

300 

200 

270 

180 

Magneto 

Magneto 

Magneto 

Magneto 

Magneto 

Magneto 

Magneto 

Splitdorf 

Splitdorf 

Bosch 

Bosch 

Bosch 

Bosch 

Bosch 

26° 

26° 

32° 

38° 

40° 

30° 

-4° 

—  4° 

—  5° 

CS 

CS 

1.5  X  CS 

1.5  X  CS 

1.5  X  CS 

1.5  X  CS 

1.5  X  CS 

cs 

CS 

IL-4R-3L- 

l!r-4R-3I^ 

1-5-3-6-2-4 

1-5-3-6-2-4 

1-5-3-6-2-4 

1-5-3-6-2-4 

1-5-3-6-2-4 

2R-4L-1R- 

2R-4L-1R- 

2I^3R 

2L-3R 

Counter 

Counter 

Counter 

Counter 

Counter 

Counter 

Clock 

clock 

clock 

clock 

clock 

clock 

2 

2 

2 

2 

2 

2 

2 

Side  of 

Side  of 

Side  of 

Cyl.  head 

Side  of 

Side  of 

Side  of 

head 

head 

head 

head 

head 

head 

18.0 

18.0 

18.0 

1.5 

1.5 

1.5 

0.021 

0.021 

0.012 

1.2  X  CS 

0.933  XCS 

0.737  X  CS 

0.5  XCS 

0.667     X  CS 

1  .  2  X  CS 

1.2       XCS. 

1.5      X  CS 

2.0  XCS 

1.894     XCS 



1.5  X  CS 

3.6 

5.95 

5.0 

14.05 

4.1 

6.19 

1.0 

1.25 

1.4 

3.30 

1.6 

2.25 

3.1 

4.40 

1.5 

1.80 

4.4 

5.62 

3.18 

6.75 

4.1 

5.65 

2.5 

3.05 

5.8 

8.92 

4.84 

9.00 

6.6 

8.70 

5.8 

8.92 

4.84 

9.00 

0.9 

0.95 

0.60 

0.881 

0.50 

0.75 

500 

500 

550 

8,768 

11,920 

13,120 

, 

3,360 

3,860 

3,612 

1,878 

2,250 

2,273 

(1)498 

(1)588 

(1)1,959 

(2,  3,  4)745 

(2,  3,  4) 

(2,  3,  4,  5,  6) 

1,192 

1,586 

(5)4,384 

(5)5,960 

(7)1,118 

78 


THE  AIRPLANE  ENGINE 


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ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


79 


The  weight  per  cubic  inch  of  piston  displacement  per  stroke  is 
an  excellent  measure  of  the  success  of  the  designer  in  keeping 
down  weight.  The  actual  horse  power  developed  depends  on 
mean  effective  pressure  and  revolutions  per  minute  and  is  affected 
by  factors  outside  the  engine  proper:  the  mean  effective  pressure 
is  determined  largely  by  the  fuel  used  and  the  carburetor  and 
manifold  resistance;  the  revolutions  per  minute  are  limited  by  the 
desirable  propeller  speed  in  ungeared  engines. 

Detailed  dimensions  of  some  of  the  most  successful  American 
and  German  engines  are  given  in  Table  4.  Weights  of  engine 
parts  are  given  in  Table  5. 

External  over-all  dimensions  for  selected  engines  are  given 
in  Table  6.  The  maximum  width  occurs  at  the  crankcase  level 

TABLE  6. — OVER-ALL  DIMENSIONS  OF  ENGINES 


Type 

Name 

Horse 
power 

No.  of 

cylinders 

Dimensions,  inches 

Maximum 

Length 

Width 

Height 

Rotary  

Radial 
air-cooled 

Vertical  

45-deg.  Vee.... 
60-deg.  Vee.  .  .  . 

90-deg.  Vee  

f  Clerget 

130 
100 
80 

170 
320 

200 
200 
160 
240 
125 
95 
240 
200 

400 

320 
360 
220 
250 
270 
400 

200 
90 
160 
180 

9 
9 
9 

7 
9 

6 
6 
6 
6 
6 
4 
6 
6 

12 

12 
12 
12 
12 
12 
12 

8 
8 
8 
8 

36.22 
24.75 
36.1 

35.7 
42.1 

63.0 
67.25 
57.08 
67.32 
63.38 
57.0 
69.88 
68.9 

69.1 

61.81 
75.98 
72.04 
55.41 
62.25 
68.3 

43.5 
50.0 
67.38 
51.3 

40.15 
37.5 
37.25 

42.2 
48.5 

22.37 
19.0 
19.92 
20.07 
18.5 
18.5 
24.09 
22.4 

26.8 

37.79 
42.52 
37.24 
35.46 
27.125 
27.9 

31.7 
30.0 
45.5 
33.5 

40.15 
37.5 
37.25 

42.2 
47.7 

39.25 
43.0 
31.88 
42.91 
41.25 
39.5 
43.62 
45.3 

43.0 

38.89 
48.03 
42.0 
33.85 
34.75 
40.1 

35.5 
27.0 
35.92 
32.7 

<  Gnome 

1  Le  Rhone        .    . 

(  ABC  Wasp  

(  ABC  Dragon  Fly  
Curtiss  K6  

Liberty  6 

Beardmore  

Galloway 

Hall-Scott  A5 

Hall-Scott  A7 

Siddeley  Puma  ...    . 

Austro-Daimler  
Liberty 

Sunbeam  Cossack  
Rolls-Royce,  Eagle  8.  . 
Rolls-Royce,  Falcon  3 
Sunbeam  Maori  
Packard  

Curtiss  K12  

Sunbeam  Arab  L  
Curtiss  OX. 

Curtiss  VX  
Hispano-Suiza  

80  THE  AIRPLANE  ENGINE 

in  vertical  engines  and  at  the  cylinder  tops  in  Vee  engines.  As 
the  tops  of  the  cylinders  are  often  above  the  fuselage,  the  width 
of  the  fuselage  is  not  necessarily  controlled  by  the  maximum 
width  of  the  engine. 

AMERICAN    ENGINES 

Liberty  Engine. — The  Liberty  engine  is  constructed  either  as  a 
six-cylinder  vertical,  or  a  12-cylinder  Vee  with  an  included  angle 
of  45  deg.  It  has  built-up  steel  cylinders,  overhead  valves  and 
camshaft,  and  battery  ignition.  The  cylinder  units  are  the  same 
in  both  constructions.  Detailed  dimensions  are  given  in  Table 
4;  weights  are  given  in  Table  5.  Longitudinal  and  transverse 
sections  of  the  12-cylinder  engine  are  shown  in  Figs.  46  and  47. 
The  performance  of  this  engine  is  shown  in  Fig.  48;  the  full 
throttle  curves  are  from  tests  with  a  dynamometer  load  and  are 
carried  up  to  2,000  r.p.m.;  the  propeller  load  curves  are  from 
tests  of  the  engine  equipped  with  its  proper  propeller  and 
mounted  on  a  torque  stand.  With  the  propeller  used  the  engine 
runs  at  about  1,700  r.p.m.  at  full  throttle  at  ground  level.  Maxi- 
mum power  is  obtained  at  about  1,850  r.p.m.  and  maximum 
economy  at  about  1,800  r.p.m.  The  brake  mean  effective  pres- 
sure, mechanical  efficiency,  and  manifold  depression  are  shown 
in  Figs.  27,  14,  and  17  respectively.  The  gear  trains  for  driving 
the  camshafts  and  various  accessories  are  shown  in  Figs.  49  and  50 
for  six-  and  12-cylinder  constructions  respectively.  Some  of  the 
details  of  this  engine  are  discussed  under  the  appropriate  headings 
in  Chapters  VI  and  VII. 

At  the  rated  speed  of  1,700  r.p.m.  the  six-cylinder  engine 
develops  232  h.p.,  and  the  12-cylinder  engine  425  h.p.  The 
diminution  in  horse  power  per  cylinder  results  from  the  lower 
mean  effective  pressure  (see  Fig.  27),  which  apparently  results 
from  lower  volumetric  efficiency.  The  weight  per  horse  power 
falls  from  2.45  Ib.  for  the  six-cylinder  to  1.99  Ib.  for  the  12- 
cylinder  engine. 

Packard. — This  engine  is  built  both  as  an  eight-  and  12-cylinder 
engine,  with  an  included  Vee  angle  in  both  cases  of  60  deg. 
It  is  very  similar  to  the  Liberty  engine  in  its  general  features  but 
has  smaller  bore  and  stroke,  a  different  method  of  cylinder 
construction  (see  Fig.  92)  and  an  underneath  carburetor  with 
induction  pipes  through  the  crankcase.  Detailed  dimensions 
are  given  in  Table  4;  weights  are  given  in  Table  5.  The  per- 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


81 


82 


THE  AIRPLANE  ENGINE 


formance  of  the  12-cy Under  engine  is  shown  in  Fig.  50.  With 
the  propeller  used  in  the  tests  the  engine  speed  is  1,600  r.p.m. 
at  full  throttle;  maximum  power  is  at  about  2,400  r.p.m.,  and 
maximum  economy  at  about  1,800  r.p.m.  The  brake  mean 
effective  pressure  and  mechanical  efficiency  are  shown  in  Figs. 
27 'and  14  respectively. 

At  the  rated  speed  of  1,600  r.p.m.  the  eight-cylinder  engine 
develops  192  h.p.,  and  the  12-cylinder  engine  280  h.p.     The 


FIG.  47. — Transverse  section  of  Liberty-12  engine. 

horse  power  per  cylinder  remains  constant.  The  weight  per 
horse  power  falls  from  2.82  Ib.  for  the  eight-cylinder  engine  to 
2.62  Ib.  for  the  12-cylinder  engine. 

Hispano-Suiza. — This  engine  is  built  in  several  sizes  as  an 
eight-cylinder  Vee  with  included  angle  of  90  deg.  In  this 
country  two  cylinder  sizes  are  built.  The  design  is  characterized 
by  steel  cylinder  sleeves  screwed  into  aluminum  water  jackets 
cast  in  blocks  of  four,  by  overhead  valves  and  camshafts,  and  by 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


83 


450 

400 
i_ 

3J    300 

1 

'     250 
o> 

1    200 
150 
100 
c    0.65 

JC 

ap  0.60 

to 
8.  0.55 

^    Q50 
0.45 

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x 

x 

X 

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full. 

/* 

Q 

Jx 

/ 

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A 

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s 

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ro^t 

-~— 

2000 


FIG. 


1200  1400  1600  1800 

Engine  Revolutions  per  Minute 

48. — Performance  curves  of  Liberty-12  engine 


Distributor 


,  Camshaft, 


Camshaft  drive, 

ikshaff  speed* 


crankshaft 
~—  speed 

Generator, 
li  crankshaft 
speed 


speed 
Side  View  Liberty  6  End  View  Liberty  |2 

FIG.  49.  FIG.  50. 

Gear  trains  of  Liberty  engines. 


84 


THE  AIRPLANE  ENGINE 


the  marine  type  of  connecting  rods.  Magneto  ignition  is  used. 
The  models  are  known  as  E  (180  h.p.)  and  H  (300  h.p.).  The  E 
engine  with  a  lower  compression  ratio  (4.72)  and  running  at 
1,450  r.p.m.  develops  150  h.p.  and  is  known  as  model  I. 

Detailed  dimensions  of  models  E  and  H  are  given  in  Table  4; 
weights  are  given  in  Table  5.  Transverse  and  longitudinal 
sections  of  model  E  are  shown  in  Figs.  51  and  52.  The  per- 
formance of  the  300-h.p.  engine  is  shown  in  Fig.  53.  With  the 
propeller  used  in  the  tests  the  engine  speed  is  1,800  r.p.m.  at  full 
throttle;  the  maximum  power  is  at  about  2,300  r.p.m.  and  maxi- 


FIG.  51. — Transverse  section  of  Hispano-Suiza  180. 

mum  economy  at  about  1,900  r.p.m.  The  brake  mean  effective 
pressure,  mechanical  efficiency,  and  manifold  depression  are 
shown  in  Figs.  27,  14  and  17,  respectively.  The  gear  trains 
for  driving  the  camshafts  and  various  accessories  are  shown  in 
Fig.  54.  Details  of  this  engine  are  described  later. 

At  the  rated  speed  of  1,800  r.p.m.  the  smaller  engine  develops 
187  h.p.,  the  larger  engine  327  h.p.  The  weight  per  horse  power 
falls  from  2.57  Ib.  for  the  smaller  to  1.94  Ib.  for  the  larger  engine. 

Wright  Engine. — This  engine  is  a  slightly  modified  Hispano- 
Suiza.  The  cylinder  jacket  is  a  little  lower,  cylinder  heads 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


85 


86 


THE  AIRPLANE  ENGINE 


thicker  and  the  lubrication  system  is  altered.     These  modifica- 
tions make  for  greater  durability. 


380 


340 


300 


260 


220 


co 


1  80 


1  40 


0.7 


0.6 


9 


yy 


§i 


x 


(Full  throttle 

— r^i — 


0  1400        1600        1800        2000       ?200      2400 

Engine  Revolutions    per  Minute 

FIG.  53. — Performance  curves  of  Hispano-Suiza  300. 


Camshaft, 
crankshaft 
speed 


Crankshaft 


Pressure  oil  pump, 
nkshaft  see 


FIG.  54. — Gear  train  of  Hispano-Suiza  180. 

Curtiss. — The  most  recent  Curtiss  models  are  the  K-6  and 
K-12;  earlier  models  include  the  OX,  VX,  V-2.  General  data  on 
these  engines  are  given  in  Table  3.  The  K-6  is  a  six-cylinder 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS  87 


88 


THE  AIRPLANE  ENGINE 


vertical  engine;  the  K-12  is  a  12-cylinder  60-deg.  Vee  engine  with 
the  same  unit  cylinder.  These  engines  are  characterized  by  the 
following  features:  cylinder  blocks  and  top  half  of  crankcase  in 
one  aluminum  casting;  cylinder  heads  in  a  separate  aluminum 
casting  bolted  to  the  cylinder  block;  steel  cylinder  liners  screwed 
into  the  cylinder  heads;  packed  watertight  joint  between  cylinder 
liners  and  aluminum  blocks;  separate  overhead  camshafts  for  the 


FIG.  56. — Transverse  section  of  Curtiss  K-12. 

inlet  and  the  exhaust  valves;  two  inlet  and  two  exhaust  valves 
per  cylinder;  sliding  cam  follower;  four-bearing  crankshaft  with 
counterweights;  crankshaft  bearings  supported  by  partition  walls 
of  upper  half  of  crankcase;  ribbed  pistons.  The  K-6  is  direct 
drive;  the  K-12  has  a  herring-bone  reduction  gear  with  a  ratio 
of  5:3,  and  an  articulated  connecting  rod. 
Longitudinal  and  transverse  sections  of  the  K-12  are  shown 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


89 


in  Figs.  55  and  56  respectively;  a  transverse  section  of  K-6  is 
shown  in  Fig.  57. 

At  the  rated  speed  of  2,250  r.p.m.  (propeller  speed  1,350  r.p.m.) 
the  K-12  develops  385  h.p.,  corresponding  to  a  mean  effective 
pressure  of  119  Ib.  per  square  inch;  at  2,550  r.p.m.  the  mean  effec- 


FIG.  57. — Transverse  section  of  Curtiss  K-6. 

tive  pressure  is  113  Ib.  per  square  inch  and  the  engine  develops 
415  h.p.  The  weight  dry  without  exhaust  manifold  is  665  Ib., 
giving  a  weight  of  1.73  Ib.  per  horse  power  at  rated  speed.  The 
K-6,  weighing  420  Ib.,  develops  approximately  200  h.p.,  which 
gives  a  weight  of  2.1  Ib.  per  horse  power.  Performance  curves 
of  the  K-12  are  given  in  Fig.  58. 


90 


THE  AIRPLANE  ENGINE 


The  OX  model  is  an  eight-cylinder  90-deg.  Vee.  It  has 
separate  cast-iron  cylinders  with  brazed  monel-metal  jackets; 
the  cylinders  in  the  two  rows  are  staggered  with  reference  to  one 
another  so  as  to  permit  the  use  of  side-by-side  connecting  rods 
on  each  of  the  four  crankpins.  The  valves  are  inclined  to  the 
cylinder  axis  and  are  operated  from  a  camshaft  inside  the  crank- 
case.  A  push  rod  operates  the  exhaust  valve  through  a  rocker 
arm;  a  pull  rod  operates  the  inlet  valve.  The  piston  is  of  alu- 
minum. The  single  carburetor  is  below  the  crankcase  and  has 
long  intake  pipes  leading  to  the  inlet  manifold. 

Longitudinal  and  transverse  sections  of  this  engine  are  shown 
in  Figs.  59  and  60.  Performance  curves  of  the  OX-5,  as  publish- 
ed by  the  builder,  are  given  in  Fig.  61. 


420 


125 


100 


300 


1100 


1200  1500  1400  1500  1600 

Revolutions  per  Minute  of  Propeller  Shaft 


FIG.  58. — Performance  curves  of  Curtiss  K-12. 


The  V-2  model  is  similar  in  general  arrangement  to  the  OX 
but  is  of  larger  size.  The  cylinders  are  of  steel  with  welded 
monel-metal  jackets;  the  valve  stems  are  parallel  to  the  cylinder 
axes.  Two  carburetors  are  used  and  are  located  at  the  level  of 
the  bottom  of  the  crankcase.  When  used  for  low  level  flying 
(up  to  6,000  ft.)  an  aluminum  liner  is  used  between  the  cylinders 
and  the  crankcase;  for  high  flights  these  shims  may  be  taken  out 
and  the  compression  ratio  thereby  increased.  A  transverse  view 
of  this  engine  is  shown  in  Fig.  62. 

Hall-Scott. — General  data  on  several  models  built  by  this 
company  are  given  in  Table  3.  The  latest  model  is  the  L-6;  it 
has  the  same  bore  and  stroke  as  the  A-5,  A-7  and  A-8  models  and 
also  the  Liberty  engine.  Longitudinal  and  transverse  sections  of 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS  91 


92 


THE  AIRPLANE  ENGINE 


FIG.  60. — Transverse  section  of  Curtiss  OXX-3. 


1100        1300        1500       1700       1900        2100 
Revolutions  per  Minute 

FIG.  61. — Performance  curves  of  Curtiss  OX-5. 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


93 


this  engine  are  shown  in  Figs.  63  and  64.  It  is  very  similar  to 
the  Liberty  6  in  its  cylinders,  valves,  pistons,  camshaft,  connec- 
ting rod  and  crankshaft.  The  crankshaft  is  supported  entirely 
by  the  upper  half  of  the  crankcase,  the  bearing  caps  being  bolted 
through  the  crankcase  by  through  bolts  which  on  the  upper  end 
act  as  cylinder  hold-down  bolts.  The  piston  pin  floats  freely  in 


FIG.  62. — Transverse  section  of  Curtiss  V-2. 


both  the  rod  and  piston.  Carburetion  is  through  two  carburetors 
and  hot-spot  water-jacketed  manifolds. 

Performance  curves  of  the  four  cylinder  L-4,  as  published  by  the 
builder,  are  given  in  Fig.  65. 

Bugatti. — The  King-Bugatti  engine,  built  in  the  United 
States,  is  a  twin-vertical  16-cylinder  engine  with  the  two  crank- 
shafts geared  to  drive  a  common  propeller  shaft.  It  is  a  modifi- 


94 


THE  AIRPLANE  ENGINE 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


95 


cation  of  the  French  Bugatti  engine.  It  has  a  number  of  special 
features  which  may  be  seen  in  the  sections,  Figs.  66  and  67.  The 
cylinders  are  of  iron,  cast  in  blocks  of  four,  with  integral  water 
jackets  except  at  the  sides,  which  are  covered  with  cast  aluminum 
plates  bolted  to  the  cast  iron.  There  are  two  inlet  valves  and  one 

exhaust  valve  per  cylinder. 
The  crankshaft  bearings  are 
supported  from  the  upper 
part  of  the  cast  aluminum 
crankcase;  the  caps,  each  of 
which  supports  two  bearings, 
extend  nearly  the  whole  width 
of  the  crankcase.  The  eight- 
throw,  nine-bearing  crankshaft 
is  in  two  halves  connected  at 
the  center  by  a  taper  and  key, 
shrunk  and  drawn  up  by  a  nut. 
Each  half  of  the  shaft  forms 
a  four-cylinder  shaft  with  the 
throws  all  in  one  plane;  the 
throws  of  the  two  sections  are 
assembled  at  right  angles.  In 


L,    120 

LJ 

E    110 


oo— -^-erfl 


90 


140 


izo  I 
no  I 
100  _« 


FIG.  64. — Transverse  section  of  Hall-Scott 
L-6. 


1000        [ZOO       1400       1600       I&OO 
Revolu+ions  per  Minute 

FIG.  65. — Performance  curves  of  Hall- 
Scott  L-4. 


assembling  the  completed  crankshafts  in  the  crankcase  they  are 
placed  in  such  relation  to  each  other  that  if  No.  8  throw  left  is  on 
top  dead  center,  No.  8  throw  right  will  be  45  deg.  past  bottom 
dead  center.  All  bearings,  including  the  connecting  rod  bearings, 
are  undercut  and  thereby  shorten  the  over-all  length  of  the  engine 
by  about  5  in.  The  valves  are  operated  by  a  single  camshaft  for 


96 


THE  AIRPLANE  ENGINE 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


97 


each  half  of  the  engine.  This  camshaft  is  operated  through 
a  bevel  gear,  driven  by  a  vertical  shaft  between  the  two  cylinder 
blocks,  which  in  turn  is  driven  by  a  bevel  gear  on  the  crankshaft. 


FIG.  67. — Transverse  section  of  Bugatti  engine. 

The  magneto  drive  (see  Fig.  245)  is  from  a  bevel  which  is  on  this 
vertical  shaft.  The  rocker  arms  are  pivoted  on  rods  running 
along  each  side  of  each  camshaft  housing.  The  cam  follower 
is  a  hardened-steel  roller;  a  smaller  roller  operates  direct  on  a 


98 


THE  AIRPLANE  ENGINE 


cap  on  top  of  the  valve  stem.     The  hollow  propeller  shaft  runs 
at  two-thirds  engine  shaft  speed. 

A  performance  curve  of  this  engine  both  with  dynamometer 
and  propeller  load  is  shown  in  Fig.  68.  Maximum  horse  power 
is  at  an  engine  speed  in  excess  of  2,400  r.p.m. 


400 


350 


300 


250 


0°, 


_  Horsepower  (Dynamometer) 
Corrected  A?  29. 92  Ha.  and 32  °F 
Observed  Horsepower(Dynamo- 
meter)  Barometer  30.45  "Hg 
Temp.  63°F. 

£_  Horsepower  (Propeller)  Corrected  for  Windage 
|         and  Faction  at2a92" and  32°F 
D-Horsepower(Propeller)Correch>dfo2332"Hg.and320F 
Observed  Horsepower  (Propeller)  Barometer  30.5" Hg. 

*>y»y      i        , 


1700         18.00        1900        2000        2100         2200        2300        2400 
Engine  Revolutions  per  Minute 

FIG.  68. — Performance  curves  of  Bugatti  engine. 

ENGLISH  ENGINES 

Rolls-Royce. — The  Rolls-Royce  engines  are  interesting  as 
probably  representing  the  highest  grade  of  design  and  manu- 
facture in  any  country.  General  dimensions  of  the  12-cylinder, 
60-deg.  Vee  "Eagle"  and  "Falcon"  models  are  given  in  the 
following  table,  which  shows  how  the  power  and  efficiency 
of  these  engines  have  been  improved  by  increasing  the 
revolutions  per  minute  and  the  compression  ratio.  Sectional 
views  are  given  in  Figs.  69  and  70.  These  engines  have  an 
epicyclic  reduction  gear  concentric  with  the  crankshaft  (see  p.  381) ; 
this  gear  is  contained  in  a  housing  bolted  to  the  front  of  the 
crankcase  and  is  not  shown  in  Fig.  69.  The  housing  for  the 
driving  gears  of  the  valve  motion,  magnetos,  and  other  accessories 
is  bolted  to  the  rear  of  the  crankcase.  The  upper  part  of  the 
crankcase  carries  the  bearings  of  the  crankshaft;  the  lower  part 
is  merely  a  deep  oil  well.  Each  cylinder  is  fastened  to  the 
crankcase  by  four  bolts.  The  engine  is  supported  by  arms  which 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


99 


100 


THE  AIRPLANE  ENGINE 


are  bolted  to  vertical  surfaces  at  the  sides  of  the  crankcase, 
thus  permitting  the  ready  adaptation  of  the  engine  to  the  airplane 
without  the  need  of  special  engine  bearers  in  the  fuselage. 


FIG.  70. — Transverse  section  of  Rolls-Royce  "Eagle." 

The  cylinders  are  of  steel  with  valve  fittings  welded  to 
short  nipples  in  the  cylinder  head  and  with  welded  jackets.  The 
aluminum  piston  is  of  special  design  without  a  skirt  at  the  middle 
third  of  its  length  (see  Fig.  98,  p.  135);  this  design  transfers 
the  gas  pressures  directly  to  the  piston  bosses  and  reduces  the 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS          101 


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102 


THE  AIRPLANE  ENGINE 


bending  stress  in  the  piston  head.     The  piston  pin  is  close  to  the 
lower  edge  of  the  piston. 

The  valves  are  of  tulip  style  (see  Fig.  115,  page  157  ),  with 
bored  stems.  Inlet  and  exhaust  valves  are  interchangeable.  A 
hardened  contact  button  is  set  in  the  end  of  the  valve  stem.  The 
crankshaft  is  provided  with  balance  weights  of  cast  steel  bolted 
to  each  crankarm;  these  have  increased  the  weight  of  the  engine 
but  have  reduced  the  pressures  on  the  main  bearings.  The 
forward  end  of  the  crankshaft  carries  a  flange  to  which  is  secured 


0.60 


220 

1400      1500      1600      HOO      1800     1900     2000 
Revolutions  per  Minute 

FIG.  71. — Performance  curves  of  Rolls-Royce  "Eagle." 

the  internally-toothed  reduction-gear  wheel  (see  Fig.  296).  A 
stud  pressed  into  the  bore  of  the  crankshaft  at  the  forward  end 
serves  as  pilot  for  the  spider  of  the  planetary  pinions.  The  con- 
necting rods  are  of  the  articulated  type,  the  secondary  rods 
working  on  pins  which  are  clamped  into  lugs  formed  on  both 
sides  of  the  main  connecting  rod. 

The  camshaft  drive  is  from  a  worm  on  the  crankshaft,  which 
is  connected  to  the  shaft  through  a  serrated  hub  joint,  and  a 
spring  coupling  clamped  by  means  of  a  disc  brake;  this  device 
protects  the  drive  against  rotary  vibration.  The  worm  drives  a 
cross  shaft  which  has  bevel  wheels  at  its  ends  driving  the  inclined 
intermediate  shafts  going  to  the  camshafts.  The  magnetos  are 
driven  from  the  cross  shaft. 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


103 


Performance  curves  for  the  " Eagle"  with  various  fuel  valve 
settings  are  given  in  Fig.  71. 

Napier  "Lion." — The  Napier  "Lion"  is  the  only  W  or  "arrow" 
type  of  airplane  engine  that  is  definitely  successful.  Details  of 


this  engine  are  shown  in  Figs.  72  and  73;  dimensions  in  Table  3. 
It  has  three  blocks,  of  four  cylinders  each,  mounted  on  a  single 
crank  case,  with  an  angle  of  60  deg.  between  the  rows.  This  engine 
is  probably  the  lightest  of  all  the  successful  water-cooled  engines; 


104 


THE  AIRPLANE  ENGINE 


it  weighs  1.86  Ib.  per  horse  power,  dry,  and  2.51  Ib.  per  horse  power 
with  its  jackets  full  of  water. 

The  cylinders  are  5j-in.  bore  and  5|-in.  stroke  which  is  an 
unusually  low  stroke-bore  ratio  and  makes  for  small  over-all  height 
and  width.  Each  block  of  cylinders  is  built  up  and  consists  of 


four  steel  liners  with  sheet-steel  water  jackets.  The  cylinders  of 
each  block  are  secured  to  an  aluminum  head  casting  to  which  they 
are  fastened  by  valve  seats,  which  pass  through  the  crown  of  each 
cylinder  and  screw  into  the  head  casting.  The  head  casting 
carries  camshafts  and  their  bearings  and  drives,  and  contains 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS  105 

the  inlet  and  exhaust  ports  and  passages.  Each  cylinder  has  two 
inlet  and  two  exhaust  valves,  the  latter  being  on  the  outside  of  the 
inclined  blocks.  The  valve  guides  are  bronze  and  a  tight  fit  in 
the  cylinder  head.  The  inlet  and  exhaust  camshafts  operate  the 
valves  directly  without  intermediate  rockers  or  plungers;  each 
valve  has  an  adjustable  tappet  head.  One  of  each  pair  of  cam- 
shafts is  driven  by  bevel  gears  on  the  inclined  and  vertical  shafts 
leading  from  the  distribution  gearing;  the  other  camshaft  is  driven 
by  spur  gearing  from  the  first  shaft.  The  crankshaft  has  four 
throws  and  is  supported  on  five  roller  bearings;  the  roller  bearings 
at  the  front  and  rear  are  fitted  direct  to  the  shaft;  the  three 
intermediate  ones  have  large  inner  races,  which  permit  the 
bearings  as  a  whole  to  be  threaded  over  the  crank  webs  and  which 
are  mounted  on  split  bushings  keyed  to  the  shaft.  The  connect- 
ing rod  assembly  consists  of  a  master  rod  and  two  side  rods 
carried  on  lugs  integral  with  the  big  end  of  the  master  rod.  The 
pistons  are  unusually  shallow,  having  a  depth  of  only  3f  in.  The 
firing  order  is : 

Propeller  End 
729 
4  11  6 

10  5  12 
1  8  3 

The  propeller  shaft  is  mounted  on  roller  bearings  with  double 
ball-bearing  thrust  block.  The  ratio  of  propeller  to  engine  speed 
is  29  to  44.  The  performance  curve  for  this  engine  is  given  in 
Fig.  74.  The  brake  horse  power  is  still  increasing  rapidly  at 
2,100  r.p.m.;  the  mean  effective  pressure  is  a  maximum  at  1,800 
r.p.m. 

Siddeley  "Puma." — This  engine  is  a  typical  six-cylinder 
vertical  engine  of  about  250  h.p.  at  1,400  r.p.m.  Cross-section 
views  are  given  in  Figs.  8  and  9;  general  dimensions  in  Table  3. 
The  cylinder  construction  consists  of  steel  liners  in  aluminum 
heads  and  jackets.  The  heads  are  cast  in  sets  of  three  and  incor- 
porate the  valve  ports  and  head  jackets.  The  barrel  jackets 
are  also  cast  in  sets  of  three  and  are  bolted  to  the  heads.  The 
steel  liners  are  screwed  cold  into  the  heads  which  are  heated  to 
about  300°C.  The  water  joint  at  the  lower  end  of  each  jacket  is 
made  by  a  screwed  gland  compressing  a  rubber  ring.  The 
three  liners  have  to  be  trued  up  by  surface  grinding  after  assembly 


106 


THE  AIRPLANE  ENGINE 


into  a  unit.  The  inlet  and  exhaust  valve  seats  are  of  phosphor 
bronze  expanded  into  the  aluminum.  The  difficulties  which 
others  have  met  in  using  this  material  at  very  high  temperatures 
have  been  overcome  by  using  an  exceptionally  hard-chilled  alloy. 
There  are  two  exhaust  valves  and  one  inlet  valve  per  cylinder. 
The*  cooling  water  enters  the  cylinder  head  at  two  places;  one  is 
in  direct  communication  with  the  water  space;  the  other  connects 
with  an  aluminum  tube  which  runs  inside  the  full  length  of  the 
jacket  and  directs  comparatively  cool  water  on  to  the  hottest 
places.  The  pistons,  connecting  rods  and  crankshaft  are  of 


FIG. 


1600    1700    1800   1900    2000  2100 
Revolutions  per  Minute 

74. —  Performance      curves 
Napier  "Lion." 


of 


1100         1300         1500        1700 
Revofu+i'ons  per  Minute 

FIG.    75.  —  Performance    curves    of 
Siddeley  "Puma." 


conventional  form.  A  roller  bearing  and  double-thrust  ball  bear- 
ing are  at  the  propeller  end  of  the  crankshaft;  the  five  intermediate 
shaft  bearings  are  supported  by  the  upper  half  of  the  crankcase. 
The  bearing  caps  are  of  aluminum  reinforced  with  webbed  plates 
of  steel;  the  caps  of  the  two  outer  journals  are  integral  with 
the  lower  half  of  the  crank  case. 

The  camshaft  is  placed  off-center  and  directly  over  the  exhaust 
valves.  It  is  made  of  steel  tubing  with  cams  pinned  in  position, 
an  arrangement  making  for  easy  and  cheap  replacement  of  worn 
cams.  The  exhaust  cams  act  directly  on  adjustable  tappet  heads 
on  the  valve  stems;  the  inlet  cams  act  on  short  drop-forged  rock- 
ers without  rollers.  The  valves  are  trumpet-shaped  and  give  good 
flow  lines. 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS  107 

A  performance  curve  for  this  engine  is  shown  in  Fig.  75; 
maximum  mean  effective  pressure  is  developed  at  about  1,000 
r.p.m. ;  maximum  horse  power  well  above  1,700  r.p.m. 

ITALIAN  ENGINES 

Fiat. — The  Fiat  650-h.p.  is  one  of  the  most  powerful  aircraft 
engines  in  use  at  present.  It  is  a  12-cylinder  60-deg.  Vee  engine 
of  unusually  large  cylinder  dimensions,  170  by  210  mm.  A 
longitudinal  section  of  this  engine  is  given  in  Fig.  76;  general 
dimensions  are  given  in  Table  3. 

The  cylinders  are  of  built-up  steel  construction  with  the 
cylinder  heads  integral  with  the  barrels  and  with  welded  water 
jackets.  Two  inlet  and  two  exhaust  valves  are  located  in  the 
head  of  each  cylinder  at  an  angle  of  25  deg.  to  the  cylinder  axis. 
The  valve  stems  work  in  phosphor-bronze  bushings.  Each 
pair  of  valves  is  closed  by  duplex  springs  located  between  the 
valve  stems  (Fig.  134).  The  camshaft  for  each  row  of  cylinders 
is  in  two  parts  which  are  driven  by  bevel  gears  from  the  center  of 
the  length  of  the  engine.  A  tubular  lay  shaft,  on  ball  bearings, 
is  mounted  in  the  center  of  the  crankcase  in  the  top  of  the  Vee  and 
is  driven  by  spur  gearing  from  the  rear  end  of  the  crankshaft; 
it  extends  only  to  the  center  of  the  engine  where  it  is  fitted  with  a 
bevel  gear  driving  the  two  inclined  shafts  which  operate  the 
camshafts.  The  connecting  rods  are  forked  type,  the  forked 
rods  being  fitted  with  bronze  bearing  shells  with  white-metal 
liners;  the  center  rod  has  a  case-hardened  steel  liner  running  on 
the  outside  of  the  bronze  shell  of  the  forked  rod.  The  crankshaft 
is  of  conventional  design;  the  main  bearings  are  held  between  the 
two  halves  of  the  crankcase,  that  is,  the  lower  half  of  the  crank- 
case  has  the  bottom  halves  of  the  bearing  housings  cast  integrally 
with  it.  The  water  pump  is  below  the  engine  at  the  middle  of  its 
length;  the  oil  pumps  are  below  the  two  ends  of  the  lower  crank- 
case.  Water  and  oil  pumps  are  driven  from  a  lay  shaft,  at  the 
bottom  of  the  lower  crankcase,  which  receives  its  motion  from  the 
crankshaft  through  spur  gears  at  the  rear  of  the  engine.  The 
induction  manifolds  are  of  sheet  copper  with  steel  flanges  brazed 
at  the  joints.  All  distribution  gears  and  driving  shafts  are 
mounted  on  ball  bearings.  There  are  four  spark  plugs  per 
cylinder  receiving  current  from  four  12-cylinder  magnetos.  The 
normal  output  of  the  engine  is  650  h.p.  at  1,500  r.p.m.;  the  maxi- 
mum is  720  b.h.p.  at  1,700  r.p.m. 


108 


THE  AIRPLANE  ENGINE 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS          109 


110 


THE  AIRPLANE  ENGINE 
GERMAN  ENGINES 


Benz. — The  Benz  230-h.p.  engine  has  six  vertical  water-cooled 
cylinders  140  by  190  mm.  Longitudinal  and  transverse  sections 
of  this  engine  are  given  in  Figs.  77  and  78;  dimensions  in  Table  4. 


FIG.  78. — Transverse  section  of  Benz  230. 

The  cylinders  are  of  cast  iron  with  pressed-steel  water  jackets 
around  the  barrels;  they  are  bolted  to  the  crankcase  by  long 
studs  which  pass  through  the  upper  crankcase  and  are  screwed 
into  the  bottom  halves  of  the  crank  bearing  housings  which  are 
cast  integral  with  the  lower  half  of^the  crankcase.  The  pistons 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


111 


are  of  cast  iron  with  the  heads  supported  by  conical  steel  forgings 
riveted  and  welded  to  the  piston  crown  and  bearing  on  the  piston 
pins  through  slots  cut  in  the  small  ends  of  the  connecting  rods; 
by  this  construction  the  greater  part  of  the  force  of  the  explosion 
is  transmitted  directly  to  the  connecting  rod.  The  connecting 
rods  are  tubular  with  internal  oil  pipes. 

Two  inlet  and  two  exhaust  valves  are  arranged  vertically  on 
each  cylinder  head  and  are  operated  through  rockers  mounted  on 
ball  bearings.  The  rockers  actuate  the  valves  through  rollers 
mounted  on  eccentric  bolts  which  permit  a  fine  adjustment  for  the 


IOOO  I200  1400  1600 

Revolutions    per  Minule 
FIG.  79. — Performance  curves  of  Benz  230.        ] 

tappet  clearance.  There  are  separate  camshafts  for  the  two 
sets  of  valves;  these  shafts  are  of  steel  tubing  and  are  inside  the 
crankcase.  They  are  driven  by  spur  gears  from  the  crankshaft 
through  an  intermediate  gear.  The  push  rods  have  hemi- 
spherical ends  at  the  bottoms  which  work  in  steel  cups  inside  the 
hollow  tappets.  The  tappets  are  slightly  offset  from  the  cam- 
shaft centers jind  carry  steel  rollers. 

The  upper  half  of  the  crankcase  has  transverse  air  passages 
cast  in  the  webs  which  serve  to  cool  the  crankcase  and  to  supply 
warm  air  to  the  carburetors.  The  lower  half  is  cooled  by  trans- 
verse aluminum  tubes,  the  air  being^led  into  one  side  by  a  large 
louvered  cowl. 

Performance  data  on  this  engine  are  given  in  Fig.  79. 


112 


THE  AIRPLANE  ENGINE 


ENGINE  DIMENSIONS  AND  ARRANGEMENTS 


113 


Maybach. — The  Maybach  300-h.p.  engine  has  six  vertical 
cylinders  built  up  of  steel  liners  screwed  into  cast-iron  cylinder 
heads  and  with  forged  steel  jackets  screwed  to  the  cylinder  heads. 
Sectional  views  of  the  engine  are  given  in  Figs.  80  and  81;  general 
dimensions  in  Table  4.  The  compression  ratio  of  this  engine, 
5.94  : 1,  is  unusually  high.  The  pistons  are  of  cast  iron.  Con- 
necting rods  are  of  square  section,  bored  out;  the  small  end  has  a 
cast-iron  floating  bushing. 

There  are  two  inlet  and 
two  exhaust  valves  work- 
ing vertically  in  cast-iron 
guides  in  each  cylinder 
head  and  operated  by 
rocker  levers  mounted  on 
roller  bearings,  each  pair 
of  valves  being  operated 
by  a  single  tappet  rod 


FIG.  81. — Transverse  section  of  Maybach  300. 


1000    1200    1400    1600    1800 
Revolutions  per  Minute 

FIG.  82. — Performance  curve 
of  Maybach  300. 


from  one  of  the  camshafts  in  the  crankcase.  The  bottom  halves 
of  the  crankshaft  bearings  are  very  deep  and  are  bolted  to  the  top 
half  of  the  crankcase.  The  lower  half  of  the  crankcase  supports 
the  oil  pumps  and  carries  detachable  oil  pumps. 

The  performance  curves,  Fig.  82,  show  maximum  mean  effec- 
tive pressure  at  1,300  r.p.m.,  maximum  power  at  1,550  r.p.m., 
and  maximum  economy  at  1,300  r.p.m. 

8 


CHAPTER  V 


MATERIALS 

The  structural  materials  available  for  airplane  engine  parts 
are  in  five  classes,  steels,  carbon  and  alloy  and  either  forged 
or  cast,  cast  iron,  aluminum  alloys,  bronze,  and  bearing  metals 
or  babbitts.  The  choice  of  metal  for  any  given  part  is  determined 
principally  by  (1)  the  suitability  of  the  metal  for  meeting  the 
stresses  and  general  operating  conditions  to  which  the  part  is 
to  be  subjected,  (2)  the  weight,  and  (3)  the  machinability 
of  the  material.  It  is  obvious  that  condition  (1)  must  always 
be  met,  but  it  is  often  possible,  especially  with  lightly  stressed 
members,  to  meet  that  condition  with  a  number  of  different 
materials;  for  example,  the  crankcase  may  be  of  cast  iron  or 
aluminum.  The  selection  as  between  these  materials  will 
usually  be  on  the  basis  of  weight  although  machinability 
or  cost  may  determine  the  final  choice.  From  the  point  of 
view  of  tensile  strength  alone  the  important  quantity  is  not 
strength  per  unit  of  cross-section  area  but  per  unit  of  weight. 
On  this  basis  the  metals  have  the  following  properties:1 


Weight  per 
cubic  foot 

Tensile 
strength, 
pounds  per 
square  inch 

Tensile 
strength  -f- 
weight  per 
cubic  foot 

Aluminum  alloy,  cast  

170 

27,000 

159 

Aluminum  alloy,  forged 

170 

60,000 

350 

Cast  iron  

480 

20,000 

42 

Gun'  metal  (bronze) 

500 

31,000 

62 

Malleable  iron  

480 

40,000 

83 

Cast  steel.  .  . 

480 

60,000 

125 

Mild  steel  

480 

60,000 

125 

High  tensile  steel 

480 

100,000 

208 

Nickel-chrome  steel  

480 

135,000 

280 

The  advantage  of  aluminum  over  cast  iron  for  castings  is  not 
only  in  the  great  reduction  in  weight  but  also  in  the  much  greater 
1  POMEROY,  Jour.  Soc.  Aut.  Eng.,  Jan.,  1920. 

114 


MATERIALS  115 


facility  for  machining.  The  substitution  of  a  lighter  material  in 
a  stressed  member  may  result  in  considerable  saving  in  weight 
even  if  the  strength-weight  ratio  of  the  lighter  material  is  not  so 
favorable  as  in  the  heavier  material.  With  castings  this  results 
from  the  fact  that  they  cannot  be  reduced  below  certain  thick- 
nesses, which  are  often  in  excess  of  strength  requirements,  because 
of  foundry  considerations.  For  machined  members,  such  as 
connecting  rods,  it  is  also  undesirable  to  go  below  certain  thick- 
nesses, and  additional  material  has  to  be  left  to  avoid  high  inten- 
sity of  stress  at  fillets  where  the  cross-section  is  changing  rapidly. 
By  using  a  lighter  metal  it  may  be  possible  to  load  all  parts  of  the 
member  up  to  the  allowable  working  stress,  and  thereby  to 
diminish  the  weight. 

An  important  design  factor  affecting  the  choice  of  metal  is  the 
fact  that  the  strength  of  members  subjected  to  bending  or  torsion 
is  proportional  to  the  cube  of  the  cross-sectional  linear  dimen- 
sions, while  the  weight  is  proportional  to  the  square  of  the  linear 
dimensions.  Thus,  comparing  two  beams  or  shafts  of  the  same 
material  and  length  and  with  similar  cross-sections  but  with 
linear  dimensions  of  1  and  1.41  respectively,  the  second  will  be 
twice  as  heavy  as  the  first,  and  (1.41)3  =  2.8  times  as  strong. 
Stiffness  is  also  an  important  consideration,  and  as  this  varies  as 
the  fourth  power  of  the  linear  cross-section  dimensions,  the 
second  beam  or  shaft  in  the  above  example  would  be  (1.41)4  =  4 
times  as  strong  as  the  first  one.  If,  in  the  second  case,  the 
allowable  stress  on  the  material  is  only  half  of  that  permissible 
in  the  first  case,  but  the  strength-weight  ratio  is  unchanged, 
the  weight  will  be  unchanged,  the  strength  increased  1.4  times 
and  the  stiffness  doubled.  For  the  same  strength  for  a  beam 
or  shaft,  with  a  stress  in  the  second  case  one-half  that  in  the  first 
case,  and  constant  strength-weight  ratio,  the  relative  linear 
dimensions  would  be  1  to  ^2,  or  1  to  1.26,  and  the  relative 
weights  1  to  (1.26)2,  or  1  to  1.588. 

Materials  for  Special  Parts. — Cylinder  liners  are  nearly  always 
of  steel.  It  is  not  necessary  to  use  a  steel  of  high  tensile  strength 
since  the  liners  cannot  be  machined  with  safety  down  to  the 
dimension  which  will  stress  the  material  fully.  The  use  of  an 
aluminum  alloy  may  be  desirable  where  short  life  is  expected  (as 
in  military  use),  but  its  resistance  to  abrasion  is  low  although  its 
heat  conduction  is  much  superior  to  that  of  steel. 

Crankshafts  are  dimensioned  for  stiffness  as  well  as  for  strength. 


116  THE  AIRPLANE  ENGINE 

Stiffness  for  given  dimensions  depends  only  on  the  modulus  of 
elasticity  and  not  at  all  upon  tensile  strength.  As  the  modulus  of 
elasticity  is  practically  constant  for  all  steels,  there  is  no  advan- 
tage, so  far  as  stiffness  is  concerned,  in  using  a  steel  of  very 
high  tensile  strength  for  shafts. 

Connecting  rods  must  be  made  as  light  as  possible  to  keep  down 
the  unbalanced  inertia  forces.  As  they  have  to  be  machined  all 
over  and  reduced  to  very  thin  sections  it  is  important  that  the 
material  should  be  readily  machined  and  free  from  flaws.  Forged 
aluminum  would  be  excellent  for  this  purpose  when  available. 

Piston  materials  are  considered  on  p.  132.  Value  materials 
are  selected  on  other  than  strength  considerations  (see  p.  159). 

Steels. — Steels  for  airplane  engine  use  should  have  the  follow- 
ing fundamental  properties  in  as  great  a  degree  as  possible:  (1) 
high  strength,  (2)  high  toughness,  (3)  great  durability,  (4) 
soundness,  (5)  ease  of  machining,  (6)  ease  of  heat  treatment,  (7) 
constancy  of  properties,  (8)  homogeneity.  The  steels  which  are 
available  are  endless  in  number  but  divide  themselves  into  certain 
types.  Those  which  are  of  most  importance  for  airplane  engine 
construction  are  listed  in  Table  7,1  which  gives  their  composition, 
heat  treatment  and  physical  properties.  Other  special  steels 
for  valves  are  discussed  on  page  160.  The  0.45  C  steel  is  an 
excellent  general-utility  steel  with  good  mechanical  strength, 
high  ductility  and  easily  machined.  The  0.9  —  1.0  C  steel 
gives  good  service  in  resisting  abrasion,  is  readily  produced  in 
sheet  form  and  does  not  need  hardening  and  tempering.  The 
case-hardening  steels  are  of  importance  for  those,  parts  which 
require  local  surface  hardening  combined  with  general  toughness; 
the  surface  hardness  of  the  0.1  C  and  the  5  per  cent  Ni  are  about 
the  same  as  shown  by  the  Brinell  test  but  the  strength  of  the 
nickel  steel  is  much  higher.  The  air-hardening  Ni-Cr  steel 
shows  the  very  high  Brinell  number  of  477;  its  use  is  principally 
for  gears.  The  3  per  cent  Ni  steel  combines  high  tensile  strength 
and  high  ductility;  the  addition  of  chromium  increases  the 
hardness.  The  chrome-vanadium  steel  has  shown  fewer  failures 
in  practice  than  most  of  the  alloy  steels,  probably  as  a  result  of 
absence  of  air-hardening  characteristics.  The  high-chromium 
or  " stainless"  steel  has  not  only  its  non-rusting  qualities  but  its 
high  mechanical  properties  to  recommend  it.  At  present  its  cost 
is  high. 

Selected  from  HATFIELD,  The  Automobile  Engineer,  June,  1920. 


MATERIALS 


117 


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118 


THE  AIRPLANE  ENGINE 


The  materials  that  have  been  employed  for  the  various  parts 
of  airplane  engines  are  given  in  Table  8,1  which  also  states  the 
nature  of  the  stresses  to  which  the  part  is  subjected  and  the  inten- 
sity of  the  service  which  it  has  to  perform.  In  addition,  there  is 
given  in  the  last  column  the  material  recommended  by  Doctor 
Hatfield  on  the  basis  of  an  unusually  wide  experience  in  investi- 
gating the  physical  properties  of  the  steels  and  the  causes  of 
failure  of  the  parts  of  airplane  engines.  There  is  naturally 
much  difference  of  opinion  among  engineers  as  to  the  best 
material  to  use  for  several  of  these  parts,  and  practice  is  by  no 
means  standardized. 

Aluminum  Alloys  are  largely  used  in  airplane  engines  on 
account  of  their  light  weight.     The  specific  gravity  of  pure  alumi- 


or 

/ 

~ensi!eStrength,Ibpe 

i-o  j£  5i  3  or 

§0  0  0  o  i 

o  o  o  §  c 

ff 

^***» 

J- 

<& 

/ 

/ 

J 

? 

S 

' 

0246      8      10     IE     14- 
Metal  AlloLjed  with  Aluminum, 
Per  Cent. 

FIG.  83. — Tensile  strength  of 
aluminum  alloys. 


14 


0       2       4       6       8       10       12 
Metal  Alloyed  vv'ith  Aluminum, 
Percent 

FIG.  84.  —  Ductility  of  aluminum 
alloys. 


num  is  2.56;  of  the  common  alloys  about  2.8;  of  steel  about  7.8. 
The  important  alloys  are  those  with  copper  and  with  zinc.  The 
effect  of  the  presence  of  these  substances  on  the  tensile  strength 

. — MATERIALS  FOR  ENGINE  PARTS 
Glossary  of  Terms  used. 


A.  H.  Ni.  Cr.  =  Air-hardening  Nickel-chro-  M.  C.  I. 

mium  Steel.  M.  C.  Steel 

=  Aluminium  Alloy.  M.  S. 

=  Case-hardening  Carbon  Ni. 

Steel.  Ni.  C.  H. 
=  Cast  Iron. 

=  Carbon  Steel.  Ni.  Cr. 
=  Chromium  Vanadium  Steel. 

=  High-carbon  Steel.  Ph.  Br 

=  12  to  14  per  cent  tungsten  Si.  Cr. 

High-speed  Steel.  Steel  C. 

=  High-tensile  Steel.  T.  Steel 


Al. 

C.  H.  C. 

C.I. 
C.  Steel 
Cr.  Van. 
H.  C.  Steel 
H.  S. 


H.  T. 


Malleable  Cast  Iron. 
Medium-carbon  Steel. 
Mild  Steel  (0.15  carbon). 
Nickel  Steel  (3  per  cent). 
Nickel  Case-hardening   Steel. 
(5  per  cent  Ni.). 
Nickel-chromium  steel  (3  per 
cent). 

Phosphor  Bronze. 
Silicon  Chromium  Steel. 
Steel  Casting. 
Tungsten  Steel. 


1  HATFIELD,  loc.  cit. 


MATERIALS 
TABLE  8  (Continued) 


119 


Part 

Nature  of 
chief  stresses 
in  service 

Intensity 
of 
service 

Materials  which 
have  been  or 
are  used 

Materials 
recommended 

Cylinder  

H'gh    temperature, 

Heavy    or 

Al.;  C.   I.;  Steel 

Al.;     80,000     Ib. 

tension  and  abra- 
sion. 

medium. 

C.;    100,000  Ib. 
C,  Steel;  Ni.;  3 
per  cent  Ni.  Cr.  ; 
Steel  Mixture. 

Steel. 

Cylinder     holding- 
down  bolts. 

Tension  and  bend- 
ing. 

Medium. 

Medium    or    M. 

S.;  Ni. 

3  per  cent  Ni. 

Cylinder  liners  

High     temperature 
and  abrasion. 

Heavy. 

C.   I.;   Steel   C.; 
Forged  Steel. 

100,000    Ib.    C. 
Steel. 

Spark-plug  body.  .  . 

High  temperature. 

Medium. 

Brass;  C.  I.;  M. 
S.;  Stainless. 

Stainless. 

Spark-plug        elec- 
trode. 

Very  high  tempera- 
ture. 

Heavy. 

T.  Steel;  Ni. 

Ni  chrome     Alloy 
or  Stainless. 

Valve  cages  

High    temperature, 
various    slight 
stresses. 

Medium.  • 

C.I. 

M.  C.  I. 

Valve-rocker     roll- 
ers. 

Abrasion. 

Light     to 
medium. 

C.  H.  C.;  Ni.  Cr. 

5    per    cent    Ni. 
C.  H.  C. 

Valves  

High    temperature, 

Heavy. 

H.  S.  Steels;  25 

Stainless. 

tension,  shock  and 
abrasion. 

per     cent     Ni.; 
Stainless  Steels; 
Ni.    Cr.;   3   per 
cent  Ni.;  H.  C. 
Steel. 

Valve  guides  

High    temperature 

Medium. 

AL;    C.    I.;    C. 

Stainless. 

and  abrasion. 

Steel;  T.  Steel. 

Valve  seats  

Abrasion,     shock 

Heavy. 

C.     I.;     M.     C. 

80,000     Ib.     C. 

and      high      tem- 
perature.            , 

Steel;    Ni.    Cr.; 
Ph.  Br. 

Steel. 

Valve  springs  

Torsion  and  bend- 

Light. 

Cr.    Van.    Steel; 

Cr.  Van.  Steel. 

ing. 

Si.     Cr.     Steel; 
H.  C.  Steel  (oil 
hardened). 

Valve  rockers  .... 

Bending,  shock  and 

Medium. 

Ni.    Cr.;  C.    H.; 

5    per    cent    Ni. 

abrasion. 

Ni.;  Bronze. 

C.  H.  ;  3  per  cent 
Ni. 

Valve-rocker   bear- 
ing. 

Abrasion  and  com- 
pression. 

Light. 

Ph.     Br.     and 
White  Metal. 

Ph.  Br. 

Water  jacket  

Very  slight. 

Light. 

Al.  ;    Copper; 

Steel  Sheet. 

Pressed      Steel; 
Sheet  Steel. 

120 


THE  AIRPLANE  ENGINE 

TABLE  8  (Continued) 


Part 

Nature  of 
chief  stresses 
in  service 

Intensity 
of 
service 

Materials  which 
have  been  or 
are  used 

Materials 
recommended 

Connecting  rod.  .  .  . 

Compression,    ten- 
sion, bending  and 
shock. 

Heavy. 

A.  H.  Ni.  Cr.; 
Cr.  Van.  Steel; 

Ni.  Cr.;  Ni. 

Ni.  Cr.;  A.  H.  Ni. 
Cr. 

Connectin  g-rod, 
big-end  bearing  .  .  . 

Abrasion  and  com- 
pression. 

Heavy. 

White  Metal;  Ph. 
Br. 

Connectin  g-rod, 
little-end  bearing.. 

Abrasion  and  com- 
pression. 

Medium. 

Ph.  Br.;  Gun 
Metal. 

Connectin  g-rod 
bolts. 

Tension  and  bend- 
ing. 

Medium. 

Ni.;  Ni.  Cr.;  Cr. 
Van.  Steel. 

3  per  cent  Ni. 

Piston  pin  

Shear,  bending  and 
abrasion. 

Heavy. 

Ni.;  160,000  to 
180,000  Ib.  C.; 
C.  H.  C.  Steel; 
C.  Steel;  Ni.; 
Ni.  Cr.;T.  Steel. 

5     per    cent    Ni. 
C.  H. 

Piston       .        ... 

Temperature,  bend- 

Heavy 

Al  •   C    I  •  Steel 

Al  •  80  000  Ib    C 

ing      and      other 

stresses. 

Drawn  or  Pres- 
sed; Ni.  Cr. 

Steel. 

Piston  ring  

High   temperature, 

Heavy 

C  I  •  H  S  Steel 

C  I 

bending       and 
abrasion. 

Bevel-gearshaft  for 
overhead     c  a  m- 
shaft. 

Torsion  and  bend- 
ing. 

Medium. 

C.  Steel;  Ni.; 
Cr.  Van.  Steel. 

5  per  cent  Ni.  C. 
H. 

Crankcase        .... 

Bending  and  vari- 

Light 

C       Steel-      Al  • 

Al 

ous  slight  stresses. 

Ni. 

Crankshaft     

Torsion,      bending, 

Heavy 

Cr     Van     Steel  • 

Ni   Cr 

and  shock. 

Ni.  Cr.;  Ni. 

Crankshaft  journal 
bearing. 

Abrasion  and  com- 
pression. 

Medium. 

Ph.  Br.;  White 
Metal. 

Propeller    reducing 
gears. 

Shear     and     bend- 
ing, abrasion  and 
shock. 

Medium. 

Ni.    C.    H.;    Ni.; 
Ni.  Cr.;Cr.  Van. 
Steel;  A.  H.  Ni. 
Cr.;  C.  H.  C. 

A.  H.  Ni.  Cr. 

Cams  

Compression    and 

Heavy. 

Ni   C   H  •  C   H 

5  per  cent  Ni.  C. 

abrasion. 

Cr. 

H. 

Cam  housings  

Negligible. 

Light. 

Al. 

Al. 

Camshaft  bearing.  . 

Abrasion  and  com- 
pression. 

Medium. 

White  Metal;  Ph. 
Br. 

White  Metal;  Ph. 
Br.;  Gun  Metal. 

Camshaft  

Torsion   and   abra- 
sion. 

Heavy. 

Ni.  Cr.  C.  H.; 
Ni.;  H.  C.  Cr. 
Steel. 

5  per  cent  Ni.  C. 
H. 

Tappets  

Compression      and 
abrasion. 

Heavy. 

C.  H.  C.  hard- 
ened on  wear- 
ing surface;  Ni\ 
Cr. 

5    per     cent     Ni. 
C.  H. 

MATERIALS  121 

of  cast-aluminum  alloy  is  shown  in  Fig.  83.  It  will  be  seen  that 
copper  increases  the  strength  up  to  about  9  per  cent,  with  7  to 
8  per  cent  of  copper  there  is  obtained  a  tough  alloy  of  a  tensile 
strength  of  20,000  Ib.  per  square  inch.  With  zinc  the  tensile 
strength  increases  up  to  about  35  per  cent;  with  this  alloy  the 
tensile  strength  reaches  50,000  Ib.  per  square  inch  but  it  is  very 
brittle  and  has  a  specific  gravity  of  3.3.  The  ductilities  of  the 
two  types  of  alloy  are  shown  in  Fig.  84.  With  both  copper  and 
zinc  present,  higher  tensile  strength  combined  with  fair  ductility 
can  be  obtained;  for  example,  with  2.75  per  cent  Cu  and  7  to 
8  per  cent  Zn  a  tensile  strength  of  28,000  Ib.  per  square  inch  and 
a  ductility  of  8  per  cent.  This  alloy  falls  off  rapidly  in  tensile 
strength  with  increase  in  temperature;  at  570°F.  the  strength  is 
9,500  Ib.  per  square  inch.  A  small  addition  of  manganese  to  a 
copper  aluminum  alloy  increases  the  strength  and  maintains  it 
better  with  increase  of  temperature. 

Forged-aluminum  alloy  is  best  represented  by  Duralumin, 
whose  composition  is  93.2  to  95.5  per  cent  aluminum,  0.5  per  cent 
magnesium,  3.5  to  5.5  per  cent  copper  and  0.5  to  0.8  per  cent 
manganese.  This  material  can  be  made  into  plates  and  tubes. 
The  tensile  strength  is  about  60,000  Ib.  per  square  inch  but 
can  be  increased  by  rolling  to  about  75,000  Ib.  per  square  inch 
though  with  loss  of  ductility;  the  elongation  of  the  alloy  is  15  to 
20  per  cent.  At  460°F.  the  tensile  strength  is  halved.  Forged- 
aluminum  is  a  good  bearing  metal. 

Both  cast-  and  forged-aluminum  alloy  have  a  modulus  of 
elasticity  of  about  10,000,000  Ib.  per  square  inch  or  about  one- 
third  that  of  steel.  The  stiffness  of  a  plate  structure  is  propor- 
tional to  the  modulus  of  elasticity  and  to  the  cube  of  its  thickness. 
An  aluminum  plate  of  the  same  weight  as  a  steel  plate  would  be 
nearly  three  times  as  thick;  its  stiffness  would  be  about  eight 
times  as  great  as  that  of  the  steel  plate. 


CHAPTER  VI 
ENGINE  DETAILS 

Cylinders. — There  is  considerable  variety  in  the  design  of 
modern  water-cooled  airplane-engine  cylinders.  In  one  impor- 
tant respect  they  are  all  in  accord,  namely,  in  the  adoption  of 
overhead  valves,  in  which  they  differ  from  the  common  auto- 
mobile engine.  They  differ  further  from  the  automobile  engine 
in  that  the  cast-iron  block  construction  is  seldom  used.  Both 
of  these  changes  have  been  made,  primarily,  in  order  to  reduce 
the  weight  of  the  engine. 

Airplane-engine  cylinders  are  formed  either  singly,  or  in 
blocks  of  two,  three  or  four.  The  single  cylinder  is  flexible  and 
permits  freedom  of  movement  of  the  cylinder  without  putting 
strains  on  other  parts  of  the  engine.  As  the  engine  is  not  held 
rigidly  it  is  desirable  to  give  all  its  parts  a  maximum  of  freedom. 
The  single-cylinder  construction  is  used  in  the  majority  of  existing 
engines  and  is  always  employed  in  those  engines  whose  cylinders 
are  all  steel.  Examples  of  this  construction  are  the  Liberty, 
Packard,  Hall-Scott  L,  Curtiss  K,  Rolls-Royce,  Napier,  Renault, 
Lorraine-Dietrich,  Fiat,  Mercedes  and  Austro-Daimler  engines. 
It  is  necessarily  employed  also  in  radial  and  rotary  engines  (see 
Chapter  VIII).  The  block  arrangement  has  a  common  jacket 
around  the  cylinders  which  may  result  in  some  slight  decrease 
in  weight  of  the  jacket  itself  but  will  usually  increase  the  weight 
of  water  in  the  jackets  and  give  inferior  water  circulation.  The 
major  advantage  of  employing  the  block  construction  is  that  it 
permits  the  cylinders  to  be  more  closely  spaced  and  thereby 
diminishes  the  over-all  length  (and  weight)  of  the  engine.  In  the 
Thomas-Morse  and  Sturtevant  engines  the  cylinders  are  in 
pairs;  in  the  Siddeley  "Puma"  in  threes;  in  the  Hispano-Suiza 
and  Bugatti  in  fours. 

Another  variation  among  airplane-engine  cylinders  is  in  the 
form  of  the  cylinder  head.  In  some  engines  (Hispano-Suiza, 
Napier)  the  cylinder  head  is  flat  and  of  the  same  diameter  as  the 
cylinder  barrel.  With  two  equal  valves  in  the  head,  the  maxi- 
mum possible  external  valve  diameter  is  half  the  cylinder  diame- 
ter minus  half  the  thickness  of  the  bridge  between  the  valves. 

122 


ENGINE  DETAILS  123 

It  is  frequently  found  that  this  diameter  is  insufficient  to  give 
the  desired  opening  for  the  admission  of  the  mixture  and  results 
in  a  low  volumetric  efficiency.  To  remedy  this  the  head  may  be 
made  with  sloping  sides  either  without  enlargement  of  diameter 
(Curtiss  OX,  Mercedes)  or  with  enlargement  (Liberty,  Lorraine 
Dietrich,  Austro-Daimler,  Fiat).  Another  common  method  of 
meeting  this  difficulty  is  by  the  use  of  multiple  valves.  The 
two  devices  may  be  used  together. 

The  most  important  factor  in  determining  the  design  of  cylin- 
ders is  the  material  employed;  the  following  constructions  are  in 
use: 

(a)  All  cast  iron,  with  cylinders  either  single  (Curtiss  OX)  or  in  blocks. 
This  construction  leads  to  excessive  weight. 

(6)  Cast-iron  barrel,  head,  and  part  of  jackets,  but  with  aluminum 
sides  to  the  jacket  (Bugatti).  This  is  for  block  construction  only.  It 
reduces  weight  and  makes  the  inside  of  the  jacket  accessible  for  machining 
and  cleaning. 

(c)  Cast-iron  barrel  and  head,  but  with  sheet  metal  jackets,  either  copper 
(Beardmore),  steel  (Benz)  or  monel-metal  (Curtiss). 

(d)  Steel  barrels,  aluminum  heads  and  jackets.     The  steel  barrel  may  be 
integral  with  its  head  (Hispano-Suiza,  Siddeley  "Puma")  or  only  a  cylin- 
drical shell  (Sturtevant).     The  aluminum  jacket  may  be  complete  (Hispano- 
Suiza,  Sturtevant)  or  may  use  the  steel  barrel  as  the  inner  wall  (Siddeley 
"Puma"). 

(e)  Steel  barrel,  cast-iron  head  and  steel  jacket  (Maybach,  Benz). 

(/)  All  steel  (Liberty,  Packard,  Hall-Scott  L,  Curtiss  K,  Rolls-Royce, 
Napier,  Renault,  Lorraine-Dietrich,  Fiat,  Austro-Daimler,  Bass6-Serve, 
Maybach).  This  construction  is  used  more  than  any  other  and  appears 
to  be  displacing  other  constructions.  The  cylinder  is  either  machined  out 
of  a  solid  forging  or  may  be  made  from  drawn  steel  tubing.  The  jackets 
are  commonly  stamped  to  shape  in  two  halves  along  a  plane  through  the 
cylinder  axis,  and  are  welded  together  and  to  the  cylinder.  The  valve 
ports  and  guides  offer  some  difficulties  as  compared  with  cast  cylinder  heads. 

Thickness  of  Cylinder  Walls. — If  p  =  maximum  gas  pressure 
in  the  cylinder,  pounds  per  square  inch,  d  —  cylinder  diameter, 
inches,  t  =  wall  thickness,  inches,  and  s  =  allowable  tensile 
stress,  pounds  per  square  inch,  then  t  =  pd/2s,  gives  the  thick- 
ness of  metal  necessary  to  withstand  the  gas  pressures.  Taking 
s  as  5,000  and  14,000  Ib.  for  cast  iron  and  forged  steel,  the  respec- 
tive wall  thicknesses  for  a  5-in.  diameter  cylinder  and  for  p  = 
500  Ib.  are  0.25  and  0.09  in.  With  cylinders  of  small  diameter 
the  thickness  may  have  to  be  increased  over  the  calculated 
values  to  ensure  sufficient  stiffness  and,  in  the  case  of  cast  iron, 


124 


THE  AIRPLANE  ENGINE 


to  offset  a  possible  lack  of  homogeneity.  No  allowance  need  be 
made  for  wear  or  reboring  as  the  engines  are  essentially  short- 
lived. 

The  clearance  space  alone  of  the  engine  is  subjected  to  the 
maximum  explosion  pressure;  the  pressures  to  which  the  walls  are 
subjected  become  progressively  less  from  the  clearance  space  to 
the  part  of  the  cylinder  at  the  lowest  point  reached  by  the  top  of 
the  piston,  below  which  point  they  become  zero.  In  addition  to 
the  gas  pressures  the  cylinder  walls  have  to  tie  the  cylinder  head 
to  the  crankcase  and  shaft  bearings  and  consequently  have  to 
withstand  the  maximum  gas  pressure  exerted  on  the  cylinder 
head.  The  thickness  of  wall  required  for  this  is  given  by 

pd       t 


The  lowest  part  of  the  cylinder  consequently  does  not  have  to  be 
more  than  one-half  the  thickness  of  the  upper  end.  In  many 
designs  the  cylinders  are  turned  of  diminishing  thickness  from 
head  to  crank  end;  in  other  designs  (Mercedes,  Lorraine  Renault, 
Liberty,  Curtiss  K)  the  upper  part  of  the  cylinder  is  reinforced  by 
stiffening  ribs  while  the  lower  part  is  without  such  stiffening. 
The  following  table  gives  some  cylinder  thicknesses  and  calcu- 
lated stresses;  a  maximum  gas  pressure  of  500  Ib.  per  square  inch 
is  assumed. 


Engine 

Diam- 
eter, 
inches 

Material 

Cylinder- 
head 
thickness, 
inches 

Cylinder- 
barrel 
thickness, 
inches 

Calculated 
maximum 
stress, 
pounds 
per  square 
inch 

Benz,  230 

5  71 

Cast  iron 

0  2600 

0  216 

6,600 

Liberty  12  

5.00 

Steel 

0.1875 

0.156 

8,000 

Curtiss,  K-12 

4  50 

Steel 

0  078 

14,400 

Hall-Scott  A5a  

5.25 

Semi-steel 

0.125 

10,500 

Austro-Daimler,  200  
Renault,  400  

5.31 

4  92 

casting 
Steel 
Steel 

0.1970 

0.138 
0.110 

9,620 
11,200 

Basse-Selve  

6.10 

Steel 

0.2700 

0.118 

12,900 

Maybach,  300  

6.50 

Steel 

0.3100 

0.110 

14,800 

The  thickness  of  the  cylinder  head  is  determined  mainly  by 
considerations  of  stiffness.  It  is  essential  that  the  valve  seats, 
which  are  located  in  the  cylinder  head,  should  be  free  from 


ENGINE  DETAILS 


125 


deformation  and  this  cannot  be  secured  unless  the  heads  are 
stiff.  In  cases  where  the  integral  cylinder  head  is  backed  by  an 
aluminum  casting  (Hispano-Suiza,  Napier)  the  thickness  cannot 
be  reduced  because  the  heat  transfer  through  the  double  thickness 
of  metal  is  poor  and  the  exhaust  valve  seat,  which  is  the  hottest 
part  of  the  cylinder,  must  have  sufficient  thickness  of  metal 
around  it  to  conduct  the  heat  away  and  prevent  warping  from 
overheating  and  unequal  expansion.  The  head  thicknesses  of  a 
few  cylinders  are  given  in  the  table  above.  Steel  cylinders  are 
generally  machined  out  of  solid  sheet  forgings,  but  in  the  case 
of  the  Liberty  engine  the  cylinders  have  been  made  from  steel 
tubing  J^  in.  thick,  whereby  a  great  reduction  is  obtained  in  the 
amount  of  metal  to  be  removed  by  machining. 

The  valve  seats  are  integral  with  the  heads  in  steel  cylinders 
but  have  to  be  cast,  or  otherwise  fastened,  into  aluminum  heads, 
With  integral  or  non-detachable  valve  seats  it  is  impossible  to 
take  out  the  valves  without  taking  off  the  cylinder  or  taking  out 
the  piston.  Detachable  valve  cages  have  been  used,  but  these 
will  always  result  in  imperfect  cooling  of  the  valve  seat.  A  com- 
promise adopted  in  the  Beardmore  and  Austro-Daimler  engines  is 
to  have  a  detachable  valve  cage  (Fig.  85) 
for  the  inlet  valve,  which  requires  little 
cooling;  the  exhaust  valve  can  then  be 
dropped  into  the  cylinder  and  withdrawn 
if  necessary  through  the  inlet  valve  seat 
opening. 

The  valve  ports  and  valve  stem  guides 
in  all  steel  constructions  are  integral  for 
each  valve  and  are  welded  to  the  cylinder 
head.  The  guides  are  provided  with  bush- 
ings of  steel  or  bronze.  The  valves  must 
work  very  freely  in  these  bushings  and  at 
the  same  time  there  must  be  no  leakage 
of  air  or  exhaust  gas  between  valve  stem  and  guide;  the  guides  are 
made  quite  long. 

Many  devices  have  been  used  for  attaching  the  cylinder  to  the 
crankcase.  The  best  method  is  to  tie  the  cylinder,  by  through 
bolts,  directly  to  the  main  bearing  caps  on  the  crankshaft  on 
each  side  of  the  cylinder.  With  two  long  bolts  on  each  side 
of  the  cylinder  tying  the  cylinder  flange  to  the  main  bearing  caps, 
the  gas  pressure  on  the  cylinder  head  is  supported  entirely 


126 


THE  AIRPLANE  ENGINE 


by  stresses  set  up  in  the  bolts  and  is  not  transmitted  to  the  crank- 
case,  which  can  consequently  be  made  lighter.  As  the  bolts  are 
parallel  to  the  cylinder  axis  and  are  symmetrically  disposed 
around  it,  this  method  of  attachment  avoids  all  distortion.  One 
pair  of  bolts  is  commonly  made  to  serve  for  two  adjacent 
cylinders  by  the  use  of  dogs  through  which  the  bolt  goes  and 

which  are  supported  equally  on 
the  flanges  of  both  cylinders 
(Fig.  80).  In  the  Bugatti 
engine  (Fig.  67)  the  through 
bolts  are  supplemented  by  studs 
in  the  crankcase. 

It  is  only  in  engines  with  a 
single  row  of  cylinders  that  the 
above  method  of  attachment  can 
be  employed.  In  Vee  and  W 
engines  other  methods  must  be 
used;  the  most  common  is  by 
studs  in  the  crankcase  which 
pass  through  the  flange  at  the 
lower  end  of  the  cylinder  (see 
Fig.  47).  In  the  Curtiss  K  en- 
gines the  steel  cylinder  is  kept 
in  place  by  the  aluminum  cyl- 
inder head,  which  is  bolted  to  an 
aluminum  jacket  cast  integral 
with  the  upper  half  of  the  crank- 
case  (Figs.  87  and  56). 

Typical  Cylinder  Designs. — 
Single  cast-iron  cylinders  are  sel- 
dom used;  a  typical  example  is 
in  the  Benz  230  engine  of  Fig. 
77.  The  barrel  and  head  are  cast  in  one  piece.  Jackets  are  of 
die-pressed  steel  welded  on  and  extending  well  down  towards 
the  base  flange.  They  are  provided  with  annular  corrugations 
to  take  care  of  expansion.  Plates  welded  in  position  in  the  jacket 
space  above  the  crown  of  each  cylinder  deflect  the  water  to  the 
exhaust  valve  pockets.  The  cylinder  barrels  extend  0.39  in. 
below  the  base  flanges  into  the  crank  chamber  and  are  held  down 
each  by  four  bolts  and  four  dogs.  The  cylinder  walls  taper 
from  6.5  mm.  at  the  top  to  5.5  mm.  at  the  base. 


FIG.  86. — Cylinder  of  Hispano-Suiza 
engine. 


ENGINE  DETAILS 


127 


An  example  of  the  block  cast-iron  construction  is  shown  in 
Figs.  66  and  67  (Bugatti).  The  cylinders  are  in  blocks  of  four. 

Composite  steel  and  aluminum  constructions  are  made  up  in 
several  ways.  In  the  Hispano-Suiza  engine  (Figs.  86  and  51) 
the  steel  liner  with  integral  head  is  threaded  throughout  the 
entire  length  of  contact  with  the  aluminum.  The  aluminum 
jacket  is  a  block  construction  for  four  cylinders  and  completely 
surrounds  the  steel  liners  so  that  valve  ports  go  through  both  the 
aluminum  and  the  steel.  For  proper  cooling  it  is  necessary  to 


FIG.  87. — Cylinder  of  Curtiss  K  engine. 

have  perfect  contact  of  the  two  metals  at  the  cylinder  head  as 
well  as  the  barrel,  otherwise  warping  of  the  steel  head  will  occur 
and  the  valve  seats  will  distort.  There  is  no  actual  contact 
anywhere  of  water  with  the  steel  cylinder.  The  steel  liners  are 
attached  to  the  crankcase  by  bolts. 

A  different  steel-aluminum  block  construction  is  employed  in 
the  Curtiss  K  engines.  In  this  case  the  steel  liners  with  integral 
heads  are  turned  with  stiffening  flanges  (Figs.  87  and  56)  on  the 
outside,  with  a  packing  retaining  flange  at  the  bottom  and  a 
central  stud  at  the  top.  The  upper  end  of  the  liner  is  of  slightly 
enlarged  diameter  and  is  threaded  on  the  outside.  The  alumi- 


128 


THE  AIRPLANE  ENGINE 


num.  cylinder  head  is  a  block  casting  into  which  the  six  cylinders 
are  screwed.  To  ensure  intimate  contact  the  threaded  stud  at 
the  center  of  the  liner  head,  which  passes  through  the  head  casting, 
is  drawn  up  by  a  nut.  The  head  casting  matches  up  with  and  is 
bolted  to  a  flange  on  the  upper  end  of  the  aluminum  cylinder 
block,  which  is  integral  with  the  upper  half  of  the  crankcase. 
Jacket  water  is  in  contact  with  the  liner  below  the  combustion 
space  only.  Water  tightness  is  obtained  by  packing  between  the 
lower  liner  flange  and  the  crankcase. 

Still  another  construction  is  used  in  the  Napier  "Lion"  engine 
(Fig.  73).  This  engine  employs  liners  with  integral  heads, 
which  are  fastened  to  a  four-cylinder  aluminum-block  head  by 
four  valve  seats  (inlet  of  bronze,  exhaust  of  steel)  in  each  cylinder. 

These  seats  are  screwed  into  the 
cylinder  head.  The  use  of  remov- 
able seats  is  in  general  objectionable 
as  it  means  less  perfect  cooling  of  the 
seats  than  is  possible  with  an  integral 
construction.  The  barrel  jackets  in 
the  latest  design  are  separate  and  of 
pressed  steel,  made  in  two  halves 
welded  together  and  to  flanges  on  the 
cylinders  at  the  top  and  bottom  of 
the  water  spaces.  In  the  design 
shown  in  Fig.  73  a  common  steel 
water  jacket  is  bolted  to  the  head 
casting  and  makes  a  joint  with  each 
cylinder  at  its  lower  end  by  means  of  a 
rubber  ring  pressed  against  a  flange  on 
the  liner  by  a  large  circular  split  nut. 

A  steel-aluminum  block  construc- 
tion in  which  the  liner  has  no  head  is 
employed    in    the   Siddeley  "Puma" 
engine    (Fig.    8).     In   this   case    the 
cast  in  sets  of  three  and  the  steel  liners 
Aluminum  water  jackets 
The  lower 


FIG.   88. — Cylinder  of  Sturte- 
vant  engine. 


aluminum  heads  are 

are  screwed  and  shrunk  into  them. 

are  also  cast  in  threes  and  are  bolted  to  the  heads. 

joint  between  the  aluminum  jacket  and  the  liner  is  made  by  a 

screwed  gland,  which  squeezes  a  rubber  ring  against  a  shoulder 

on  the  outside  of  the  liner.     The  valve  seats  are  of  phosphor 

bronze  expanded  into  the  aluminum  head. 


ENGINE  DETAILS 


129 


FIG.  89. — Cylinder  of  Maybach  engine. 


FIG.  90.— Cylinder  of  Benz  300  engine. 


130 


THE  AIRPLANE  ENGINE 


Another  construction  in  which  a  headless  steel  liner  is  em- 
ployed is  the  Sturtevant  engine,  Fig.  88.  The  aluminum  cylin- 
ders are  cast  in  pairs  and  are  provided  with  closely  fitting  steel 
liners;  the  heads  are  also  of  aluminum  with  inserted  valve  seats 
and  are  in  pairs.  A  peculiarity  of  this  construction  is  that 
the  aluminum  cylinders  are  bolted  direct  to  the  crankcase;  the 
steel  liner  does  not  transmit  any  longitudinal  forces  and  is 
perfectly  free  to  expand. 

A  composite  steel  and  cast-iron  con- 
struction is  used  in  the  Maybach  engine 
(Figs.  89  and  81).  The  steel  barrel  is 
screwed  with  a  buttress  thread  into  a 
cast-iron  head  and  comes  up  against  a 
soft  brass  washer  which  prevents  water 
leakage.  The  water  jacket  is  a  machined 
forging  and  screws  on  to  the  cylinder  head 
(see  details,  Fig.  89)  where  it  is  sweated  in 
position  with  soft  solder;  the  lower  joint  is 
kept  watertight  by  a  gland  and  rubber 
ring.  The  valve-stem  guides  have  cast- 
iron  bushings. 

In  the  Benz  300  engine  (Fig.  90)  the 
cylinder  head  is  of  cast  iron  and  the  rest 
of  the  cylinder  is  of  steel.  There  are  ports 
for  two  inlet  valves  and  one  exhaust  valve. 
The  steel  liner  screws  into  the  cast-iron 
head  and  makes  a  watertight  joint  by  bed- 
ding into  cement  in  the  small  groove  into 
which  the  top  of  the  liner  goes. 

The  all-steel  construction  is  exemplified  in  the  Liberty  cylinder 
(Fig.  91).  This  design  has  a  bumped  head  and  obliquely  set 
valves.  The  jackets  and  valve  ports  are  welded  to  the  sheet 
cylinder.  Details  of  this  welding  are  shown  for  the  very  similar 
Packard  engine,  Fig.  92. 

In  the  Austro-Daimler  engine  (Fig.  93)  the  cylinders  are 
all  steel  with  pressed  steel  jackets  and  twin  inlet  and  exhaust 
valves  in  the  cylinder  heads.  The  valve  pockets  are  welded  in 
position  with  the  exception  of  one  inlet  valve  (Fig.  85),  which  is 
detachable  with  its  seat  and  guide.  The  cylinders  taper  from 
4  mm.  at  the  top  to  3  mm.  at  the  middle  and  increase  again  to 
4  mm.  at  the  bottom;  the  jackets  are  1  mm.  thick.  The  bottom 


FIG.    91.  —  Cylinder    of 
Liberty  engine. 


ENGINE  DETAILS 
W 


131 


W: 


w 


FIG.  92. — Cylinder  head  of  Packard  engine  showing  welding,  W,  of  the  water 

jacket. 


P 


FIG.  93. — Cylinder  of  Austro-Daimler  engine. 


132 


THE  AIRPLANE  ENGINE 


of  each  jacket  is  flanged  over  and  welded  to  a  bevelled  flange 
machined  on  the  barrel. 

Pistons.— One  of  the  most  important  steps  in  the  improve- 
ment of  the  airplane  engine  has  been  the  general  substitution  of 
an  aluminum  alloy  for  the  cast  iron  that,  until  very  recently,  was 
universally  employed  for  the  piston  material.  The  piston  must 
be  as  light  as  possible  in  order  to  keep  down  inertia  forces  and  at 
the  same  time  must  be  thick  enough  in  the  crown  to  conduct  the 
heat  away  rapidly  from  the  center  to  the  circumference,  where  it  is 
taken  up  by  the  cylinder  walls.  With  cast-iron  pistons,  reduc- 
tion in  weight  has  resulted  in  many  troubles  and  especially  in  the 
burning  and  cracking  of  the  piston  head.  Measurements  by 
Hopkinson  in  a  Siddeley  engine  show  the  cast-iron  piston  head 
has  a  temperature  of  over  900°F.  This  is  borne  out  by  similar 
measurements  on  the  pistons  of  air-cooled  cylinders  by  Gibson. 


FIG.  94. — Aluminum  piston. 


FIG.  95. — Cast-iron  piston. 


With  aluminum  pistons  this  temperature  is  reduced  to  about 
400°F.  and  at  the  same  time  the  piston  is  considerably  lighter. 
The  two  pistons  of  Figs.  94  and  95  for  a  100  by  140  mm.  air- 
cooled  engine  have  weights  of  1.26  Ib.  and  1.77  lb.,  or  a  reduction 
in  weight  of  29  per  cent  by  the  substitution  of  aluminum  alloy 
for  cast  iron;  these  pistons  show  the  temperature  difference 
noted  above. 

The  lower  temperature  of  the  aluminum  piston  has  other 
important  results.  It  reduces  the  rise  in  temperature  of  the 
incoming  charge  during  the  suction  stroke  and  thereby  increases 
the  volumetric  efficiency  of  the  engine,  and  it  permits  the  use  of  a 
higher  compression  ratio  without  danger'  of  preignition  and 
thereby  increases  both  the  capacity  and  efficiency  of  the  engine. 
The  horse  power  of  an  engine  can  be  increased  at  least  5  per  cent 
by  the  use  of  aluminum  pistons. 

The  piston  should  be  designed  with  heat  dissipation  in  mind 
as  much  as  the  other  piston  functions.  The  heat  received  at  the 
center  of  the  head  must  pass  out  radially  to  the  circumference,  and 


ENGINE  DETAILS 


133 


this  should  be  provided  for  either  by  adequate  thickness  of  metal 
from  center  to  periphery  or  by  the  provision  of  radial  ribs  which 
serve  the  double  function  of  heat  carriers  and  stiffening  members. 
Furthermore,  the  thickness  of  the  cylindrical  wall  must  not  be 
cut  down  behind  the  first  ring  as  the  heat  has  to  flow  downward 
from  the  crown. 

The  friction  between  the  piston  and  the  cylinder  is  by  far  the 
largest  item  of  mechanical  loss  (see  p.  24)  and  should  be  kept 
as  low  as  possible.  Its  high  value  results  apparently  from  the 
partial  carbonization  of  the  lubricant  clinging  to  the  walls  under 
the  action  of  the  high  gas  temperatures  and  of  the  slight  but 
unavoidable  leakage  of  burning  gases  past  the  piston  rings.  As  a 
result,  the  viscosity  of  the  lubricant  is  greatly  increased.  The 


FIG.  96. — Slipper  piston. 

extent  of  the  friction  depends  upon:  (1)  The  pressure  of  the 
piston  against  the  cylinder  walls,  which  governs  the  thickness  of 
the  film;  the  friction  appears  to  be  proportional  to  the  average 
loading;  (2)  the  area  of  the  bearing  surface;  (3)  the  quantity  of 
the  lubricant  on  the  walls;  the  friction  increases  with  this  quan- 
tity; (4)  the  temperature  of  the  walls,  which  controls  the  viscosity 
of  the  lubricant.  Of  the  means  adopted  to  reduce  this  friction 
loss  the  most  prominent  is  the  cutting  away  of  the  piston  skirt  on 
the  sides  which  do  not  support  side  thrust. 

This  practice  has  the  further  advantage  of  reducing  the  weight 
of  the  piston.  An  example  of  such  a  piston  is  shown  in  Fig.  96. 
It  will  be  seen  that  the  piston-pin  bosses  in  this  design  are 
supported  by  vertical  transverse  ribs,  which  pass  within  a  distance 
from  the  center  of  the  head  of  a  little  more  than  half  of  the  radius 
of  the  piston.  This  is  an  important  feature  in  keeping  the  piston 
head  cool;  the  heat  absorbed  by  the  central  portion  of  the  head 


134 


THE  AIRPLANE  ENGINE 


can  pass  down  these  ribs  to  the  gudgeon-pin  bosses  and  to  the 
skirt.  This  in  turn  permits  the  use  of  a  thinner  crown,  especially 
as  the  stiffness  of  the  crown  is  greatly  increased  by  the  support 
of  the  ribs.  The  thickness  of  the  lubricant  film  is  diminished 
by  providing  holes  in  the  slippers  through  which  excess  oil  is 
squeezed  out.  If  both  slippers  are  designed  for  the  same  inten- 
sity of  pressure,  then  areas  of  the  two  slippers  will  be  different. 
Such  a  design  is  shown  in  Fig.  97;  the  supporting  ribs  are  no 
longer  parallel. 


FIG.  97. — Piston  with  unequal  slippers. 

One  of  the  important  troubles  with  pistons  is  the  slap  which 
occurs  when  the  side  thrust  is  transferred  from  one  side  to  the 
other.  The  amount  of  this  slap  depends  on  the  clearance 
between  the  piston  and  cylinder.  The  cold  clearance  must  be 
larger  with  aluminum  than  with  cast  iron  as  the  expansion  is 
much  greater.  The  pistons  of  Figs.  94  and  95  have  cold  clear- 
ances of  0.026  and  0.020  in.  respectively;  the  hot  clearances  for 
both  are  0.008  in.  The  clearances  should  be  greatest  at  the  top 
and  where  the  temperatures  are  highest  and  should  diminish  as 
the  bottom  of  the  skirt  is  approached.  With  aluminum  pistons  a 


ENGINE  DETAILS 


135 


cold  clearance  over  the  top  lands  of  about  0.005  in.  per  inch 
diameter  is  necessary.  Such  a  piston  will  be  noisy  when  cold. 
For  cast  iron  the  cold  clearance  should  be  0.003  in.  per  inch 
diameter  at  the  top,  and  0.00075  at  the  base.  The  clearances  for 
special  engines  are  given  in  Table  4. 

In  order  to  prevent  piston  slap,  or  the  opposite  danger  of 
seizing  when  hot,  the  practice  has  arisen  of  insulating  the  piston 
skirt  from  the  ring-carrying  portion  of  the  piston.  This  is  most 
readily  accomplished  by  the  use  of  a  piston  with  piston-pin  bosses 
carried  by  ribs  (Fig.  97)  and  with  the 
skirt  or  slippers  separate  from  the 
upper  portion  of  the  piston.  A  design 
of  this  character  is  shown  in  Fig.  98 ;  in 
this  case  a  complete  skirt  is  used.  Such 
constructions  are  satisfactory  only  for 
cylinder  diameters  up  to  about  5  in.; 
for  larger  sizes  the  piston  cooling  will  be 
inadequate.  The  clearance  necessary 
for  the  skirts  of  aluminum  divided  slip- 
per pistons  is  about  the  same  as  that  re- 
quired for  the  normal  cast-iron  piston. 

Another  method  of  reducing  piston  slap  is  by  offsetting  the 
wristpin  by  a  small  amount,  usually  not  more  than  J4  m-  The 
object  of  this  construction  is  to  cause  the  piston  to  tilt  slightly 
about  the  piston  pin  and  therefore  to  pass  progressively  instead 
of  abruptly  from  one  cylinder  wall  to  the  other.  Such  offset 
is  shown  in  Figs.  96  and  97. 

The  composition  of  the  aluminum  alloy  used  in  German  pistons 
is  given  below.  These  pistons  are  usually  die  castings  and  have  a 
tensile  strength  of  28,000  to  31,000  Ib.  per  square  inch  and 
extension  of  4^  per  cent  as  against  about  one-half  those  quan- 
tities for  sand  castings. 


FIG.  98.— Divided-skirt 
piston. 


1 

g 

£ 

S3 

I 

N 

i—  i  • 

Silicon 

a 

H 

2 

I 

i 

Alumin 

Benz  230  h  p          

6  02 

12  13 

1  42 

0.31 

0 

0 

Tr. 

Tr. 

80.  12 

Austro-Daimler  200  h.p  

7.67 

1.33 

1.32 

0.52 

2.21 

0 

Tr. 

0.29 

86.66 

Basse-Selve  270  h.p  

1.90 

15.62 

1.06 

0.45 

0 

0 

0 

o 

80.97 

136 


THE  AIRPLANE  ENGINE 


Piston  weights  (including  piston  rings  and  gudgeon  pin) 
vary  from  0.19  Ib.  (Austro-Daimler)  to  0.25  Ib.  (Liberty)  per 
square  inch  of  piston  area  in  aluminum  construction;  for  cast 
iron  the  weight  may  exceed  0.42  Ib.  (May bach). 

Typical  Pistons. — Figure  99  is  the  May  bach  cast-iron  piston. 
Figure  100  is  the  Benz  cast-iron  piston  with  a  thin  crown  and  with 


__^ 


FIG.  99. — Maybach  cast-iron  piston. 

a  hollow  conical  steel  pillar  riveted  to  the  piston  head  and  resting 
directly  on  the  middle  of  the  piston  pin.  The  small  end  of  the 
connecting  rod  is  cut  away  to  avoid  interference  with  this  pillar. 
With  this  arrangement  the  gas  pressure  is  transmitted  in  the  line 
of  the  connecting  rod  and  there  is  no  bending  moment  on  the 
wrist  pin.  The  ribbed  aluminum  construction  is  shown  in  Fig. 


FIG.   100.— Benz  230  cast-iron  piston. 

101  for  the  same  engine.  The  top  clearances  for  the  two  pistons 
are  0.02  and  0.03  in.  respectively;  the  bottom  clearances  are 
0.004  and  0.014  in.  respectively.  The  weights  are  6.72  Ib.  and 
4.90  Ib.  respectively,  complete  with  rings  and  setscrews  but 
without  piston  pins;  the  saving  in  weight  is  27  per  cent.  The 
ribless  construction  is  typified  by  the  Liberty  engine  piston  shown 
in  Fig.  102;  oil  grooves  are  provided  on  the  piston  skirt. 


ENGINE  DETAILS 


137 


Piston  rings  are  of  dense  gray  cast  iron,  fully  machine-finished, 
peened  on  the  inner  curved  surface  and  exactly  ground  to  size 
upon  the  outer  curved  surface.  With  very  narrow  rings  semi- 
steel  is  used.  The  rings  may  be  either  of  the  concentric  or 
eccentric  types.  The  ends  are  commonly  chamfered  at  an  angle 
of  30  to  40  deg.  but  stepped  ends  are  also  used ;  the  gap  when  in 
the  cylinder  is  about  ^40  the  diameter  of  the  piston.  Three 


FIG.  101. — Benz  230  aluminum  ribbed  piston. 

rings  are  commonly  used,  but  four  rings  are  found  occasionally. 
Two  narrow  rings  are  sometimes  used  in  one  groove.  A  scraper 
ring  near  the  bottom  of  the  skirt  is  sometimes  used  to  clear  excess 
oil  from  the  cylinder  walls;  the  same  result  is  obtained  by  the 
use  of  perforations  through  the  skirt.  In  some  pistons  (Figs. 
99  and  100)  the  lowest  ring  acts  as  a  scraper  and  has  a  groove 
below  it  through  which  small  holes  are  drilled  to  the  interior  of  the 
piston  to  drain  away  any  excess  of  oil. 


138 


THE  AIRPLANE  ENGINE 


Piston  or  Gudgeon  Pin. — The  piston  pin  is  usually  of  steel, 
machined  to  size,  case-hardened  and  ground.  It  is  always 
hollow.  It  is  most  commonly  fully  floating,  that  is,  it  has  bearing 
in  the  end  of  the  connecting  rod  as  well  as  in  the  piston  bosses, 
with  some  end  motion  as  well.  The  pin  will  then  rotate  and 

local  wear  will  be  avoided.  The 
bushing  in  the  connecting  rod 
is  also  floating  in  many  engines. 
On  the  cold  motor  the  pin 
should  be  a  mild  driving  fit  in 
the  bosses  and  a  running  fit  in 
the  connecting-rod  bushing. 
When  warm  the  aluminum 
bosses  expand  more  than  the 
bushing  and  the  pin  becomes 
free.  With  standard  piston 
types,  as  in  Figs.  99-102,  the 
piston  pin  is  comparatively  long 
and  is  subjected  to  considerable 
bending  stress.  By  the  adoption 
of  the  slipper  piston  (Figs.  96 
and  97)  or  the  divided  skirt 
(Fig.  98)  the  bosses  are  brought 
closer  together  and  the  pin  is 
shortened.  It  can  consequently 
be  made  of  smaller  diameter  and 
lighter  and  if  fully  floating  will 
show  no  wear. 

Connecting  Rods. — In  vertical 
engines  the  connecting  rods  are  of 
uniform  (non-tapering)  circular 
or  I-section,  with  solid  small 
ends  and  marine  type  big  ends.  An  example  of  the  circular 
section  (Benz  230)  is  shown  in  Fig.  103;  the  I-section,  assembled 
with  the  piston  (Siddeley  "Puma"),  in  Fig.  104.  The  small  end  is 
usually  provided  with  a  bronze  bushing,  although  in  the  Maybach 
engine  this  is  replaced  by  a  perforated  cast-iron  floating  shell 
0.124  in.  thick.  The  large  end  has  a  babbitted-bronze  shell. 
The  tubular  rods  used  in  several  German  engines  are  sometimes 
provided  (Fig.  103)  with  a  centered  internal  pipe  for  lubricating 
the  small  end.  The  Benz  rod  has  a  number  of  radial  holes 


FIG. 


102. — Liberty  ribless  aluminum 
piston. 


ENGINE  DETAILS 


139 


drilled  in  the  big  end  to  reduce  weight  and  has  the  top  of  the 
small  end  cut  away  to  permit  direct  application  of  the  gas 
pressure  load  on  the  crankpin  through  a  pillared  piston  as  in 
Fig.  100. 

In  Vee  engines  several  connecting  rod  arrangements  are  used. 
In  the  Curtiss  OX  and  V,  Sturtevant  and  Thomas-Morse  engines 


FIG.   103. — Benz  tubular  connecting  rod. 

the  cylinders  in  the  two  rows  of  the  Vee  are  staggered  so  that  the 
connecting  rods  do  not  lie  in  the  same  plane.  With  this  arrange- 
ment the  big  ends  of  each  pair  of  cylinders  lie  side  by  side  on  the 
same  crankpin.  Such  an  arrangement  results  in  an  increase  in 
the  over-all  length  of  the  engine. 


140 


THE  AIRPLANE  ENGINE 


With  the  cylinders  of  each  pair  opposite-  one  another,  as  in 
the  general  practice  with  Vee  engines,  the  connecting  rods  are  in 
the  same  plane  and  special  arrangements  must  be  made  to  connect 
them  both  to  the  crankpin.  Two  arrangements  are  in  use,  the 
forked  rod  and  the  articulated  rod.  The  forked  rod  is  used  in  the 
Liberty,  Hispano-Suiza,  Packard,  and  Fiat  engines.  Each  pair 
of  rods  consists  of  a  plain  rod  and  a  forked  rod.  The  forked  rod 
clamps  the  big  end  bronze  bushing;  the  plain  rod  works  on  the 


FIG.   104. — Siddeley  "Puma"  I-section  connecting  rod. 

outside  of  this  bronze  bushing  between  the  two  forks  of  the  forked 
rod.  The  rods  for  the  Liberty  engine  are  shown  in  Fig.  105.  The 
bearing  is  prevented  from  rotating  in  the  forked  rods  by  dowel 
pins.  In  the  Fiat  engine  the  bottom  ends  of  the  fork  are  fastened 
together. 

In  the  articulated  rod  assembly  a  master  rod  is  used  and  a 
short  rod  is  attached  to  a  pin  which  is  held  in  the  upper  half  of 
the  big  end  of  the  master  rod.  Ordinarily  the  master  rods  are  all 
placed  on  one  side  of  the  Vee,  but  occasionally  (Renault)  the 
master  rods  alternate  with  short  rods.  A  good  example  of  a 


ENGINE  DETAILS 

o 


141 


/A 


FIG.  105. — Forked  connecting  rods  of  Liberty  engine. 


FIG.  106.— Master  rod  of  Benz  300  engine. 


142 


THE  AIRPLANE  ENGINE 


master  rod  is  shown  in  Fig.  106  for  the  Benz  300,  60-deg.  Vee 
engine.  The  rod  is  tubular;  the  pin  for  the  small  rod  is  held  by 
two  clamp  screws.  In  the  Renault  engine  (Fig.  107)  the  pin  for 
the  small  rod  is  of  the  same  diameter  as  the 
piston  pin  so  that  both  ends  of  the  small  rod 
are  alike. 

In  W  engines  the  articulated  rod  is  most 
common.  The  Napier  "Lion"  uses  a  central 
master  rod,  on  each  side  of  which  is  mounted 
an  articulated  rod  (Fig.  108)  carried  on  pins 
fixed  in  lugs  integral  with  the  big  end  of  the 
master  rod.  The  main  rod  is  of  I-section;  the 
side  rods  are  tubular  and  carry  bronze  bush- 
ings at  both  ends.  Each  side-rod  pin  is 
tapered  at  one  end,  fits  into  a  tapered  hole  in 
the  corresponding  lug,  and  is  drawn  up  tight  by  a  bolt  screwing 
into  the  pin  at  the  taper  end. 

In  the  Lorraine  W  engines  the  two  outer  rods  bear  on  the 
cylindrical  outer  surface  of  the  big  end  of  the  master  rod,  the 


FIG.  107. — Articu- 
lated connecting  rods 
of  Renault  engine. 


FIG.  108. — Articulated  connecting  rods  of  Napier  "Lion." 

bearing  slippers  covering  less  than  half  the  circumference.  Two 
circular  steel  rings  hold  the  two  halves  of  the  big  end  of  the 
master  rod  together  and  hold  the  slippers  of  the  outer  rods  to  the 
outer  surface  of  the  master  rod. 


ENGINE  DETAILS  143 

The  articulated  arrangement  of  connecting  rods  suffers  from 
some  disadvantages  as  compared  with  an  arrangement  in  which 
the  connecting  rods  are  always  radial  to  the  crankpin.  The 
short  rod  is  materially  shorter  than  the  master  rod — usually  at 
least  20  per  cent  shorter — and  consequently  causes  greater 
angularity  of  that  rod  and  increased  side  thrust  in  the  cylinder. 
The  explosion  pressures  transmitted  along  the  short  rod  do  not 
act  directly  on  the  crankpin  but  impose  stresses  on  the  master 
rod  which  under  unfavorable  conditions  may  be  serious.  For 
example,  with  a  90-deg.  Vee 
and  with  explosion  pressure 
reached  30  deg.  before  dead 
center  in  the  short-rod  cylinder, 
the  force  acting  on  the  master 
rod  CD  (Fig.  109)  would  be 
directed  along  the  line  AE. 
The  reaction  of  this  force  at 
C  can  be  readily  found  and, 
treating  the  master  rod  as  a  FIG.  109. — Diagram  of  articulated  con- 
cantilever  loaded  with  this  re-  necting  rods< 
action  force  at  C  and  held  on  the  crankpin,  the  stress  at  any 
section  of  the  rod  can  be  ascertained. 

Connecting  rods  up  to  the  present  time  have  always  been 
made  of  steel.  The  use  of  forged  aluminum  rods  would  mate- 
rially reduce  the  weight  of  the  reciprocating  parts  and  the  bearing 
pressure  at  the  big  end. 

Crankshafts. — Crankshafts  are  made  of  alloy  steel  (nickel 
or  chrome-vanadium)  and  are  usually  forged  in  one  piece.  An 
exception  to  this  is  the  Bugatti  eight-throw  shaft  which  is  made 
in  two  lengths.  In  airplane  practice  there  are  usually  bearings  on 
both  sides  of  each  throw;  this  gives  great  stiffness  to  a  light  shaft. 
The  arrangement  common  in  automobile  engines  of  two  or  more 
throws  between  main  bearings  has  been  employed  by  the  Sturte- 
vant,  Thomas-Morse  and  Duesenberg  engines  and  is  still  used 
in  the  Curtiss  K  engines  (see  Fig.  55).  The  long  crankarms 
(between  the  first  and  second,  and  between  the  fifth  and  sixth 
cranks)  of  this  engine  have  centers  of  gravity  which  do  not 
coincide  with  the  axis  of  the  shaft  (as  in  four-cylinder  engines) 
and  consequently  produce  an  unbalanced  moment  about  the 
crank  axis.  This  can  be  balanced  by  a  counterbalance  weight 
between  the  two  center  crankpins  which  are  in  line.  The  addi- 


144  THE  AIRPLANE  ENGINE 

tion  of  such  a  counterbalance  weight  sets  up  an  undesirable 
bending  moment  on  the  long  central  crankpin  and  also  puts 
more  load  on  the  two  central  main  bearings,  which  have  to  resist 
the  moments  created  by  these  unbalanced  masses.  To  eliminate 
these  objectionable  conditions  it  is  desirable  to  balance  directly 
the  masses  of  the  two  long  crankarms.  This  is  accomplished  in 
the  Curtiss  K  engines  (Fig.  55)  by  applying  balance  weights 
directly  to  the  long  crankarms.  To  obtain  the  greatest  possible 
balancing  effect  with  the  least  weight,  aluminum  spacers  are 
inserted  between  the  steel  balance  weights  and  the  crankshaft, 
the  balance  weights  being  held  to  the  crankshaft  by  steel  bolts. 

The  crankshaft  and  pins  are  always  made  hollow  and  holes 
are  drilled  through  the  crank  webs  for  oil  passages  connecting  the 
hollow  crankpins  and  journals.  The  open  ends  of  the  shaft  and 
crankpins  are  plugged  by  screw  plugs  (Hispano-Suiza)  or  by 
discs  or  caps  which  are  expanded  or  brazed  into  place,  or  in  some 
cases  (Liberty,  Siddeley  "Puma")  are  held  in  position  by  bolts 
which  tie  together  a  pair  of  caps.  In  this  last  case  the  bolts  can 
be  used  to  obtain  rotational  balance,  the  method  being  to  use 
special  bolts  thickened  in  the  middle.  The  caps  in  different  con- 
structions are  of  duralumin,  gun  metal,  steel  and  other  metals 
and  may  be  used  with  or  without  gaskets. 

Main  bearings  and  crankpin  bearings  are  nearly  always  of 
babbitted  bronze.  Occasionally  a  ball  bearing  is  used  at  one 
end;  at  the  rear  end  in  the  Hispano-Suiza;  at  the  front  end  in  the 
Fiat.  In  the  Napier  "Lion"  with  three  cylinders  on  each  crankpin 
and  with  a  heavy  big  end,  the  main-bearing  pressures  are  very 
high  and  roller  bearings  are  used  throughout.  Double-thrust 
ball  bearings  are  usual  and  are  placed  just  behind  the  propeller 
hub. 

The  pressure  on  main  bearings  is  high  and  demands  con- 
siderable oil  circulation,  not  only  for  lubrication  but  also  for 
cooling.  The  babbitt  tends  to  flake  off  unless  the  bronze  has  been 
tinned  before  casting  the  babbitt,  in  which  case  a  perfect  bond  can 
be  obtained;  mechanical  holding  of  the  babbitt  by  holes,  dove- 
tails or  screw  threads  is  generally  found  unsatisfactory. 

The  bearing  pressure  in  a  six-throw,  seven-bearing  crankshaft 
is  greatest  at  the  center  main  bearings  because  the  crank  throws 
on  the  two  sides  of  it  are  in  line  so  that  the  dynamic  loads  imposed 
on  the  two  cranks  are  in  phase.  Consequently  the  center  main 
bearing  is  often  made  longer  than  the  intermediate  bearings  to 


ENGINE  DETAILS  145 

diminish  the  intensity  of  pressure.  In  Rolls-Royce  and  Fiat 
engines  the  center  bearings  are  60  per  cent  longer  than  the 
intermediate  bearings.  In  the  Liberty  engine  the  maximum 
load  on  the  center  main  bearing  is  7,700  Ib.  or  1,675  Ib.  per 
square  inch  of  projected  area;  the  mean  unit  bearing  pressure  is 
1,265  Ib.  per  square  inch.  As  the  rubbing  velocity  is  19.5  ft. 
per  second,  the  friction  work  is  F  =  f  X  19.5  X  1,265  ft.-lb. 
per  second,  where  /  is  the  coefficient  of  friction.  On  the  inter- 
mediate main  bearings  of  this  engine  the  maximum  load  figures 
out  as  7,250  Ib.  or  1,580  Ib.  per  square  inch  of  projected  area;  the 
mean  unit  bearing  pressure  is  700  Ib.  per  square  inch.  The  end 
main  bearings  receive  loading  on  one  side  only  and  show  a  maxi- 
mum load  of  4,025  Ib.  or  a  maximum  unit  pressure  of  815  Ib. 
per  square  inch  and  a  mean  unit  pressure  of  610  Ib.  per  square 
inch. 

The  crankpin  bearing  pressure  for  the  Liberty  engine  has  a 
maximum  value  of  4,980  Ib.  or  932  Ib.  per  square  inch  of  projected 
area;  the  mean  unit  bearing  pressure  is  642  Ib.  per  square  inch. 
The  crankpin  pressures  used  in  the  German  engines  are  somewhat 
lower,  ranging  from  a  mean  unit  bearing  pressure  of  402  Ib.  per 
square  inch  in  the  Austro-Daimler  to  585  Ib.  per  square  inch 
in  the  Maybach.  The  crankpin  pressure  is  mainly  due  to 
inertia  and  centrifugal  forces — the  gas  pressures  have  com- 
paratively little  effect.  This  is  evidenced  by  the  fact  that  the 
wear  on  crankpins  bearings  is  on  the  side  remote  from  the  piston, 
that  is,  on  the  side  subjected  only  to  inertia  and  centrifugal  forces. 

Crankshafts  are  subjected  to  stresses  which  vary  rapidly 
in  sign  and  magnitude  and  consequently  are  especially  liable  to 
fail  from  fatigue  of  material.  The  weakest  point  is  generally  at 
some  place  where  there  is  a  sudden  change  in  cross-section  and 
poor  distribution  of  stress.  It  is  particularly  important  that  the 
fillets  at  the  junctions  of  the  crankpins  and  journals  with 
the  crank  webs  should  be  of  adequate  size.  Tests  to  determine  the 
desirable  size  of  the  fillet  have  been  conducted  recently  in 
England;  they  show  that  the  steel  is  materially  weakened  if  the 
fillet  is  less  than  %  in.  radius. 

In  the  discussion  of  torque  on  page  47  it  has  been  shown  that 
the  maximum  torque  at  the  propeller  end  of  the  crankshaft  of  a 
six-cylinder  engine  is  actually  less  than  the  maximum  value  at 
the  rear  crankpin.  Consequently  there  is  no  need  for  any 
increase  in  diameter  of  the  crankshaft  from  rear  to  front  in  that 
10 


146  THE  AIRPLANE  ENGINE 

case.  The  free  end  of  the  crankshaft  is  subjected  to  much  more 
severe  conditions  than  the  propeller  end.  The  free  end  is,  as  it 
were,  wound  up  when  the  maximum  torque  is  applied  to  it  and 
released  when  the  torque  diminishes.  At  certain  speeds  this 
alternate  winding  up  and  release  may  coincide  with  a  natural 
period  of  vibration  of  the  crankshaft,  and  in  that  case  the  shaft 
will  vibrate  excessively  and  the  reciprocating  masses  attached 
to  it  will  also  vibrate  and  impart  their  vibration  to  the  whole 
structure.  Such  torsional  vibration  could  be  reduced  by  the  use 
of  a  flywheel  on  the  free  end.  It  has  given  much  trouble  in 
various  six-cylinder  engines  and  has  been  largely  responsible  for 
the  failure  of  engines  with  eight  crank  throws. 

With  six-cylinder  engines  of  200  to  300  h.p.  the  freely  vibrating 
shaft  has  a  frequency  which  is  usually  about  6,000  vibrations  per 
minute;  for  four-cylinder  engines  this  frequency  is  higher,  and  in 
single-crank  radial  or  rotary  engines  it  may  be  as  high  as  20,000. 
The  period  of  vibration  can  be  determined  by  striking  a  series 
of  light  blows  at  regular  intervals;  the  vibrations  will  increase 
markedly  when  the  frequency  of  the  blows  coincides  with  the 
natural  period  of  the  shaft.  In  a  six-cylinder  engine  the  impulses 
are  three  per  revolution  so  that  the  dangerous  speed  for  a  shaft 
with  vibration  frequency  is  6,000  per  minute  of  6,000  -f-  3  = 
2,000  r.p.m.  The  next  most  dangerous  speed  would  be  1,000 
r.p.m.  With  eight-cylinder  engines  the  dangerous  speeds  would 
be  J^  and  M  the  vibration  frequency. 

The  addition  of  counterweights  to  the  crankshaft  as  a  means 
of  obtaining  rotary  balance  is  of  value  mainly  in  reducing  bearing 
pressures.  The  centrifugal  force  arising  from  unbalanced 
rotating  weights  acts  radially  from  the  center  and  produces  con- 
siderable bearing  pressures.  Under  airplane-engine  conditions  a 
lower  total  weight  is  obtained  by  omitting  counterbalances  and 
giving  the  bearing  sufficient  area  and  stiffness  to  support  the 
centrifugal  forces.  With  higher  speeds  of  rotation  the  need  for 
counterbalance  weights  increases. 

Crankshafts  are  drop  forgings  and  are  usually  made  in  dies 
when  the  quantity  warrants  it;  in  other  cases  they  are  cut  from 
large  billets.  The  dies  may  be  made  of  cast  iron  when  the 
number  of  forgings  required  is  small;  for  quantity  production 
they  are  of  steel. 

Strength  of  Shafts. — Shafts  should  be  designed  for  their 
strength  in  shear.  For  a  solid  circular  shaft  of  diameter  d  in. 


ENGINE  DETAILS 


147 


subjected  to  a  bending  moment  M  in  pound-inches  and  a  torsion 
T,  also  in  pound-inches,  and  with  a  maximum  permissible  intensity 
of  shearing  stress  at  the  outer  surface  of  the  shaft  of  /  Ib.  per 
square  inch, 

d3  =  5.1 


For  a  hollow  shaft  of  outside  diameter  d2  and  inside  diameter 
the  equation  becomes 


d* 


=  5.1 


FIG.  110. — Propeller  hub  of  Hispano-Suiza  engine. 

With  M  =  24,500,  T  =  54,000,  and  /  =  16,000,  these  equations 
give  d  =  2%  in.  and  if  d2  is  assumed  to  be  3  in.  di  is  2.275  in. 
Under  these  conditions  the  hollow  shaft  will  weigh  56  per  cent 
as  much  as  the  solid  shaft.  A  hollow  shaft  of  still  larger  outside 
diameter  would  be  lighter  and  stiffer  but  would  require  larger 
and  heavier  bearings  and  would  result  in  increased  rubbing 
velocities  at  the  bearings  and  increased  friction. 

Propeller  hubs  in  American  practice  are  mounted  on  a  tapered 
extension  of  the  crankshaft.  The  Hispano-Suiza  hub  (Fig.  110) 
is  a  good  example  of  standard  practice.  It  is  keyed  to  the 
engine  shaft,  which  is  given  a  taper  of  1  in  10,  and  is  threaded  at 
the  end  to  receive  a  long  nut  which  is  used  for  forcing  the  hub 
on  the  taper.  The  inner  flange  is  integral  with  the  tapered  hub; 
the  outer  flange  has  splines  which  fit  in  grooves  on  the  outer  end 


148 


THE  AIRPLANE  ENGINE 


of  the  hub  and  permit  axial  movement  of  about  1  in.  for  adjust- 
ment to  the  thickness  of  the  propeller,  which  is  held  between  the 
two  flanges.  Eight  bolts  hold  the  flange  and  the  propeller 
together.  Rotation  of  the  hub  on  the  engine  shaft  is  prevented 
by  a  key.  The  long  nut  is  held  in  position  by  a  locknut  or  a 
locking  pin. 

The  Benz  engine  (Fig.  Ill)  employs  a  hub  which  is  bolted  to  a 
flange  at  the  end  of  the  crankshaft  and  has  an  outer  flange  which 
fits  on  the  splined  end  of  the  hub. 

Crankcases. — The  crankcase  has  to  serve  several  functions: 
it  has  to  tie  various  parts  of  the  engine  together;  it  has  to  with- 
stand stresses  due  to  gas  pressures,  and  bending  moments  due  to 
unbalanced  forces;  it  contains  the  lubricating  oil;  it  supports 


FIG.   111. — Propeller  hub  of  Benz  engine. 

various  auxiliaries;  and  it  has  to  support  the  engine  as  a  whole. 
The  stresses  in  the  crankcase  are  chiefly  of  the  two  kinds  sug- 
gested above.  The  explosion  pressure,  acting  on  the  cylinder 
head,  puts  the  cylinder  under  tension  and  this  should  be  supported 
as  directly  as  possible  by  connecting  members  from  the  cylin- 
der to  the  corresponding  lower  main  crankshaft  bearings.  In 
vertical  engines  this  can  be  accomplished  very  satisfactorily  by 
the  use  of  through  bolts  from  the  lower  cylinder  flange  to  the 
lower  bearings  (see  Fig.  112),  using  transverse  webs  in  the 
upper  crankcase  as  distance  pieces.  In  Vee  and  W  engines 
this  construction  is  not  possible  and  it  is  necessary  to  fasten  the 
cylinders  and  the  lower  bearings  to  the  upper  crankcase,  and 
transmit  the  tension  through  the  transverse  webs  to  the  lower 
bearings.  The  lower  half  of  the  crankcase  is  sometimes  cast  as  a 
unit  with  the  lower  bearings  but  this  practice  has  nothing  to 
recommend  it.  The  assembly  of  cylinders  and  upper  crankcase 
should  sustain  all  the  stresses  due  to  gas  pressures. 


ENGINE  DETAILS 


149 


The  unbalanced  centrifugal  and  inertia  forces  acting  through 
the  main  bearings  subject  the  crankcase  to  bending  moments 
which  change  continuously  in  direction  and  magnitude.  It  is 
necessary  that  the  crankcase  should  have  sufficient  stiffness  to 
withstand  these  bending  moments  without  objectionable  deflec- 
tions. For  this  purpose  a  webbed  box  structure  has  been  found 
most  satisfactory.  It  is  easily  possible  to  obtain  the  necessary 
stiffness  by  utilizing  the  upper  crankcase  only  as  a  stressed 
member.  The  lower  crankcase  is  preferably  used  only  as  an  oil 


FIG.   112. — Transverse  section  of  crankcase. 

container  and  a  support  for  oil  pumps  and  other  auxiliaries. 
This  practice  can  be  seen  in  the  Hall-Scott  L  (Fig.  64),  Bugatti 
(Fig.  67),  Curtiss  K  and  OX  (Figs.  55  and  59),  Napier  "Lion" 
(Fig.  72),  Maybach  (Fig.  81),  and  Benz  engines  (Fig.  78).  In 
other  engines  the  lower  crankcase  carries  also  the  lower  halves  of 
the  end  main  bearings,  as  in  the  Hispano-Suiza  (Fig.  51)  and  Sid- 
deley  "Puma"  engines  (Fig.  9).  These  arrangements  permit  of 
easy  accessibility.  In  the  Liberty  engine  (Fig.  47)  the  lower 
crankcase  has  transverse  webs  and  the  crank  chamber  is  divided 
into  six  separate  chambers;  a  similar  construction  with  double 
transverse  webs  is  employed  in  the  Fiat  engine  (Fig.  76). 


150  THE  AIRPLANE  ENGINE 

The  transverse  webs  are  often  cut  away  in  places  for  lightness. 
Aluminum  alloy  is  universally  used  for  crankcases. 

The  lower  crankcase  serves  as  an  oil  sump.  The  earlier 
practice  of  keeping  a  considerable  body  of  oil  (wet  sump)  in  the 
bottom  of  the  crankcase  is  now  being  superseded  by  the  dry  sump 
which  is.  kept  drained  by  a  scavenger  pump  or  pumps.  Wet  sumps 
are  shown  in  the  Benz  engine  (Fig.  78),  Hispano-Suiza  (Fig.  51), 
Hall-Scott  L  (Fig.  64),  and  Curtiss  OX  engines  (Fig.  59).  In 
the  last  case  the  sump  is  separated  by  drainage  plates  from  the 
rest  of  the  crankcase.  The  advantage  of  the  dry  sump  is  in 
avoiding  drowning  the  cylinder  with  oil  in  case  the  engine 
operates  momentarily  upside  down  or  in  any  posture  approximat- 
ing that  position.  The  drainage  point  of  the  sump  is  usually 
in  the  middle,  but  in  some  cases  the  scavenger  pumps  take  oil 
from  both  ends  (Liberty,  Napier  "Lion7')-  Oil  cooling  is  carried 
out  in  the  lower  crankcase  in  the  Austro-Daimler  engine  by 
casting  outside  cooling  ribs  running  longitudinally  along  the 
bottom  of  the  crankcase  and  attaching  a  sheet  of  aluminum  in 
such  a  way  as  to  form  an  air  duct  along  the  whole  underside  of  the 
engine.  In  the  Basse-Serve  engine  an  oil  cooler  with  air  tubes  is 
fastened  to  the  bottom  of  the  crankcase  but  has  no  direct  com- 
munication with  the  inside.  In  the  Curtiss  K  engine  (Fig.  55) 
oil  cooling  is  effected  inside  the  crankcase  by  the  jacket  water  on 
its  way  to  the  pump;  this  arrangement  serves  also  to  heat  up  the 
oil  quickly  after  starting  the  engine  and  puts  the  lubrication 
system  into  normal  operation  earlier  than  would  otherwise  be 
possible. 

German  airplane  engines  usually  have  provision  for  air  cooling 
of  the  crankcase.  In  the  Benz  engine  (Fig.  77)  the  support- 
ing webs  for  six  of  the  main  bearings  form  air  passages  trans- 
versely across  the  engine.  Two  of  these  serve  as  air  intake 
passages  to  the  two  carburetors,  which  are  thereby  supplied  with 
heated  air.  In  addition  the  lower  crankcase  is  traversed  by  18 
aluminum  tubes,  30  mm.  in  diameter;  air  is  scooped  into  the 
tubes  through  an  aluminum  louvered  cowl  on  one  side  of  the  engine 
and  discharged  through  a  reversed  cowl  on  the  other  side. 


CHAPTER  VII 
VALVES  AND  VALVE  GEARS 

Location  of  Valves.  —  The  diversity  of  valve  locations  which 
characterizes  automobile  practice  is  not  found  in  modern  airplane 
engines.  L-head  and  T-head  arrangements  have  been  supplanted 
by  overhead  valves  which  permit  the  simplest  form  of  cylinder 
head  and  watex  jacket,  shortest  and  most  direct  passage  of  the 
gases,  and,  with  overhead  camshafts,  a  considerable  simplifica- 
tion in  the  valve  gear  and  a  reduction  in  the  number  of  rubbing 
or  contact  points.  The  valves  may  either  be  seated  in  a  flat 
head  in  which  case  the  stems  are  parallel  to  the  cylinder  axis  or 
the  head  may  be  domed,  in  which  case  the  valve  stems  are 
inclined  to  the  cylinder  axis. 

Valve  Lift.  —  Valves  are  always  of  the  poppet  type  with  bevelled 
seats.  They  are  opened  by  cams  which  operate  either  directly 
or  indirectly;  they  are  closed  by  springs.  The  valve  (Fig.  113) 
has  a  face  which  is  usually  about  25  per 
cent  wider  than  the  seat  on  which  it 
closes  and  a  stem  which  passes  through 
a  long  guide  (often  provided  with  a 
bushing)  and  which  connects  with  the 
valve  by  a  rounded  fillet.  The  bevel 
of  the  valve  face  is  usually  45  deg.  but 
30  deg.  is  sometimes  used;  The  width 
of  the  valve  face  must  be  small  to 
ensure  gas  tightness  and  is  usually 
about  one-fourth  the  lift  of  the  valve. 
The  width  of  the  valve  seat  is  usually 
less  than  0.1  in.  The  free  area  for 
the  passage  of  the  gas  through  the 
fully  opened  valve  may  be  taken  approximately  as  irdh  where 
d  is  the  smallest  diameter  of  the  bevelled  valve  face  and  h 
is  the  lift.  This  area  should  not  be  greater  than  the  free  open- 
ing through  the  valve  seat.  Neglecting  the  area  occupied 


FIG. 


1 13.  —  Typical  airplane 
engine  valve. 


by  the  valve  stem,  irdh  =    d2,  or  h 

151 


gives  the  lift  which 


152  THE  AIRPLANE  ENGINE 

makes  the  gas  passage  area  equal  to  the  free  opening  through  the 
valve;  usually  h  varies  from  one-fifth  to  one-sixth  of  the  outside 
valve  diameter.  Values  will  be  found  in  Table  4. 

With  a  valve  lift  of  one-quarter  of  its  diameter  the  gas  flow 
for  a  given  pressure  drop  is  found  by  experiment  to  be  about  67 
per  cent  of  the  flow  through  an  unobstructed  port1;  with  a  lift 
of  one-half  the  diameter  this  is  increased  to  from  80  to  90  per  cent. 
These  "  coefficients  of  efflux"  are  found  to  be  the  same  for  all 
pressure  drops,  and  for  valves  of  different  sizes  at  equal  lifts 
expressed  in  per  cent  of  their  respective  diameters.  The  experi- 
ments were  carried  out  with  continuous  flow  which  presumably 
would  give  results  differing  from  those  actually  occurring  under 
the  operating  conditions  of  intermittent  flow.  The  earlier  in- 
vestigations of  Lucke2  indicate  coefficients  of  efflux  lower  than 
those  given  above  ;  the  variation  with  lift  is  probably  of  the  right 
order  of  magnitude. 

The  volume  of  gas  passing  the  inlet  valve  per  unit  of  time  is 
approximately  equal  to  the  piston  displacement  in  that  time;  the 
volume  of  gas  passing  the  exhaust  valve  is  from  two  to  three 
times  as  great.  If  the  mean  piston  speed  during  the  suction 
stroke  is  s  in.  per  second  and  the  piston  diameter  is  D  in.  the 
mean  gas  velocity  V  in  feet  per  second  past  the  valve  is  given  by 

~D2-s  =  12V<jrdh 
4 

or 

ft.  per  second 


where  d  and  h  are  in  inches.  This  velocity  is  approximate,  as  the 
equation  assumes  the  cylinder  to  fill  with  a  mixture  at  atmos- 
pheric pressure  and  temperature.  As  the  piston  speed  is  varying 
throughout  the  stroke,  the  gas  velocity  will  vary  and  will  have  a 
maximum  value  which  is  nearly  twice  the  mean  value. 
For  the  Liberty  engine  the  mean  gas  velocity  is  189  ft.  per  second; 
for  most  engines  it  varies  from  about  150  to  200  ft.  per  second. 
The  pressure  drop  past  the  valve  to  obtain  this  velocity  can 
be  obtained  with  sufficient  accuracy  from  the  equation  V2  =  2gh, 
where  h  is  the  pressure  drop  measured  in  feet  of  air.  To  convert 
this  to  a  pressure  drop,  -i,  measured  in  inches  of  water,  using  the 

1  LEWIS  and  NUTTING,  4th  Ann.  Report,  Nat.  Adv.  Comm.  Aeronautics,  1918. 

2  Trans.  A.  S.  M.  E.t  1905,  Vol.  27. 


VALVES  AND  VALVE  GEARS  153 

conversion  factor,  1  in.  of  water   =  5.2  Ib.  per  square  foot,  we  have 
h  =  5.2  X  i  X  v 

where  v  is  the  volume  of  1  Ib.  of  the  explosive  mixture.     The 
quantity  v  is  obtained  from  the  perfect  gas  equation 
v  =  RT/p  =  52  T/p. 

At  ordinary  atmospheric  pressure  and  temperature,  v  has  a  value 
of  about  13. 

V2   =2gX  5.2  X  13  Xi 
and 

.         F2 
4,350 

For  gas  velocities  from  150  to  200  ft.  per  second  the  corresponding 
pressure  drops  are  from  5.2  to  9.2  in.  of  water. 

The  pressure  drop  through  the  valve  is  important  as  affecting 
the  pressure  in  the  cylinder  at  the  end  of  the  suction  period  and, 
thereby,  the  volumetric  efficiency.     This  pressure  is  controlled 
also  by  the  frictional  resistances  to  the  flow  through  carburetor, 
manifold  and  gas  ports,  by  the  time  and  rate  of  valve  opening 
and  closing,  and  by  other  factors.     At  midstroke,  when  the  piston 
velocity  is  a  maximum,  the  gas  velocity  past  the  valve  will  be  a 
maximum;  if  its  value  is  twice  the  mean  velocity  the  pressure 
drop  will  be  four  times  the  mean  value.     In  the  Liberty  engine 
the  pressure  drop  corresponding  to  a  mean  velocity  of  189  ft.  per 
second  is  8.2  in.  of  water;  the  pressure  drop  past  the  valve  at 
midstroke  would  then  be  about  4  X  8.2  =  33  in.  of  water.     Fric- 
tional resistances  will  increase  this  quantity.     During  the  latter 
half  of  the  stroke  the  quantity  of  gas  passing  the  valve  will  be 
greater  than  in  the  first  half  on  account  of  the  existence  of  this 
greater  vacuum.     That  is,  during  the  first  half  of  the  stroke  a 
vacuum  is  being  created  in  the  cylinder;  during  the  last  half  this 
vacuum  is  being  filled  up.     Near  the  end  of  the  stroke  the  valve 
is  closing  which  cuts  down  the  area  for  gas  flow  and  limits  the 
filling  up  process.     In  most  engines  (see  p.  175)  the  final  closure 
of  the  valve  does  not  occur  till  after  the  dead  center  and  the 
additional  time  so  obtained  for  the  admission  of  the  charge  has 
important  results  in  increasing  the  weight  of  charge  admitted. 
The  continued  flow  of  gas  into  the  cylinder  past  the  dead  center 
position  is  due  not  only  to  the  existence  of  a  vacuum  there  but 
also  to  the  inertia  of  the  column  of  gas  in  the  induction  system. 
It  is  not  desirable  to  delay  the  final  closure  of  the  valve  long 


154 


THE  AIRPLANE  ENGINE 


enough  to  establish  atmospheric  pressure  in  the  cylinder;  this 
procedure  would  not  increase  the  volumetric  efficiency  because 
of  the  diminished  volume  of  the  charge  resulting  from  the 
return  of  the  piston.  The  actual  closing  of  the  valve  is  some- 
times delayed  till  the  piston  has  returned  10  per  cent  of  its  stroke 
so  that  the  volume  of  gas  at  the  beginning  of  compression  is 
correspondingly  reduced  and  the  compression  ratio  is  less  than 
the  cylinder  dimensions  would  indicate. 

The  gas  flow  area  past  the  valve  has  been  seen  to  be  propor- 
tional to  dh,  or  since  h  is  proportional  to  d,  the  area  is  proportional 
to  d2.  The  larger  this  area  the  lower  will  be  the  gas  velocity, 
and  the  less  the  frictional  resistances  and  the  pressure  drop. 


FIG.   114. — Valve  arrangements. 

With  flat-headed  cylinders  and  with  no  increase  in  diameter 
in  the  combustion  space  the  valve  diameter  must  be  considerably 
less  than  half  the  cylinder  diameter;  in  the  Hispano-Suiza  180  it 
is  42  per  cent,  in  the  Hispano-Suiza  300  it  is  40  per  cent.  With 
enlarged  heads  the  ratio  of  valve  diameter  to  cylinder  diameter 
can  easily  be  made  50  per  cent  as  in  Liberty  engines,  or  even 
55  per  cent  as  in  the  Curtiss  VX  engine. 

To  obtain  a  larger  valve  area,  dual  valves  may  be  employed. 
The  Benz  200  has  dual  valves  of  1.693  in.  diameter  as  against  one 
valve  of  2.205  in.  diameter  in  the  Hispano-Suiza  300  engine  of  the 
same  cylinder  diameter;  the  possible  gain  in  valve  area  is  17^ 
per  cent.  A  comparison  is  shown  in  Fig.  114  of  the  maximum 
valve  areas  for  the  Hispano-Suiza  300  engine,  a  with  single 
valves,  b  with  dual  intake  and  dual  exhaust  valves  and  c  with  dual 
intake  and  single  exhaust.  The  single  valves  are  of  the  same 


VALVES  AND  VALVE  GEARS  155 

diameter  as  in  the  actual  engine;  the  width  of  bridge  between 
valves  is  kept  constant  and  the  clearance  from  the  cylinder  walls 
is  the  same  as  in  the  actual  engine.  The  maximum  dual  valve 
area  in  b  is  19  per  cent  and  in  c  is  45  per  cent  greater  than  the 
single  valve  area  in  a.  In  Fig.  114  d  the  single  inlet  valve 
is  shown  of  the  same  area  as  the  dual  exhaust  valves;  with  this 
arrangement  both  inlet  and  exhaust  valve  areas  are  increased  23 
per  cent  as  compared  with  a.  The  increase  in  valve  area  obtain- 
able by  the  use  of  dual  valves  becomes  greater  as  the  cylinder  dia- 
meter increases  since  it  is  not  necessary  to  change  the  width  of 
bridge  between  valves  or  the  cylinder  wall  clearance  with  change 
in  cylinder  diameter.  For  example,  with  a  cylinder  diameter  of 
7  in.  and  bridge  width  0.69  in.  and  wall  clearance  0.206  in.  the 
use  of  dual  valves  increases  the  area  24  per  cent  as  compared 
with  19  per  cent  gain  in  the  Hispano-Suiza  300  with  cylinder 
diameter  5.51  in.  and  the  same  bridge  width  and  clearance. 

A  further  development  along  the  same  lines  is  the  use  of 
triple  inlet  and  exhaust  valves  as  in  the  800-h.p.  Sunbeam  Sikh 
engine.  This  arrangement  leads  to  a  decrease  in  valve  area 
except  for  large  cylinders.  With  a  7-in.  cylinder  and  the  same 
bridge  width  and  clearance  as  above  the  decrease  in  valve  area 
with  triple  valves  in  line  (Fig.  114  e)  is  35  per  cent  as  compared 
with  dual  valves;  with  the  valves  arranged  in  a  circle  (Fig.  114/) 
the  decrease  in  valve  area  is  16  per  cent.  The  latter  arrange- 
ment would  require  a  more  complicated  valve  gear  than  is 
necessary  with  the  three  valves  in  line.  With  larger  cylinder 
and  smaller  bridge  width  the  comparative  showing  of  the  triple 
valves  would  improve. 

The  comparisons  just  presented  between  multiple  and  single 
valves  are  based  on  the  assumption  of  the  same  lift  expressed  in 
per  cent  of  diameter.  If  dual  valves  are  used  and  the  actual  lift 
is  kept  the  same  as  for  the  corresponding  single  valve,  the  gas 
flow  through  the  multiple  valves  will  be  increased  by  about  20 
per  cent  over  the  figures  given;  for  example,  the  dual  valves  of 
Fig.  114  b  will  increase  the  gas  flow  about  40  per  cent  above  that 
through  the  single  valve  for  a  given  pressure  drop.  The  diam- 
eters of  the  dual  valves  can  be  reduced,  without  decreasing  the 
gas  flow  below  that  through  a  single  valve,  by  maintaining  the 
same  lift  as  for  a  single  valve.  For  example,  a  2.5-in.  valve  with 
25  per  cent  lift  (0.625  in.)  gives  the  same  flow  as  two  1.75-in. 
valves  with  25  per  cent  lift  (0.44  in.)  or  two  1.5-in.  valves  with 


156  THE  AIRPLANE  ENGINE 

0.625-in.  lift  (42  per  cent).  This  last  arrangement  has  certain 
advantages;  the  area  of  valve  presented  to  gas  pressure  is  only 
72  per  cent  of  the  area  of  the  single  valve  and  the  weight  of  the 
two  valves  will  be  only  56  per  cent  of  the  weight  of  the  single 
valve,  assuming  the  weights  to  vary  as  d2-5.  With  valve  spring 
tensions  proportional  to  valve  weights  the  power  required  to 
operate  these  dual  valves  will  be  less  than  half  the  power  required 
to  open  the  single  valve.  Other  advantages  are  that  a  larger 
proportion  of  the  cylinder  head  can  be  jacketed  because  it  is  not 
occupied  by  the  valves;  that  the  valve  cooling  will  be  better 
because  the  circumference  of  the  two  valves  is  20  per  cent  greater 
than  that  of  the  single  valve  and  the  distance  travelled  by  the 
heat  is  only  60  per  cent  as  great;  and  the  distortion  of  the  valve 
will  be  less  in  consequence  both  of  lower  temperature  and  smaller 
diameter. 

The  area  of  the  exhaust  valve  opening  affects  the  volumetric 
efficiency  since  it  determines  the  pressure  of  the  residual  gases 
in  the  combustion  space  at  the  time  of  opening  of  the  inlet  valve 
and  determines  the  back  pressure  on  the  piston  during  the  exhaust 
stroke.  The  main  problem  of  the  exhaust  valve  is  heat  dissipa- 
tion. The  exhaust  valve  is  heated  during  the  explosion  stroke 
and  by  the  exhaust  gases  as  they  pass  out.  The  greater  the 
velocity  of  the  gases  the  greater  is  the  heating  of  the  valve.  Heat 
abstraction  from  the  valve  is  principally  by  conduction  to  the 
seat  and  thence  to  the  jacket  water,  but  is  also  by  conduction  to 
the  stem  and  thence  to  the  guide.  The  inlet  valve  gives  no 
trouble  from  this  source  as  it  is  cooled  by  the  entering  charge. 
Increase  of  exhaust  valve  area  by  increasing  its  diameter  is 
objectionable  because  it  results  in  increased  temperature  of  the 
valve.  The  heat  received  by  the  valve  is  approximately  pro- 
portional to  the  square  of  its  diameter  while  the  area  of  the  heat- 
abstracting  seat  is  proportional  to  the  first  power  of  the  diameter. 
Furthermore,  as  valve  diameter  increases  the  mean  distance  the 
heat  has  to  travel  increases  and  this  results  in  increased  valve 
temperature  and  also  in  valve  warping.  The  maximum  practica- 
ble size  of  exhaust  valve  seems  to  be  about  3  in.  diameter.  A 
limit  to  valve  size  is  also  set  by  the  valve  weight,  which  increases 
about  as  d2-5  while  the  area  increases  as  d2.  As  the  valves  are 
closed  by  springs,  the  dimensions  of  the  springs  have  to  increase 
rapidly  with  increase  of  valve  diameter  and  lift  in  order  to  give 
sufficiently  quick  closure.  The  desirable  method  of  increasing 


VALVES  AND  VALVE  GEARS 


157 


exhaust  valve  area  is  therefore  by  the  use  of  multiple  valves  and 
not  by  increase  in  diameter  of  a  single  valve.  The  arrangement  of 
Fig.  1 146  with  dual  exhaust  valves  and  single  inlet  valve,  is  often 
employed  (Siddeley  "Puma/7  Benz  300,  ABC  Dragon-fly), 
while  the  reverse  arrangement  of  two  inlets  and  one  exhaust 
valve  is  seldom  used  (Bugatti).  Ordinarily,  if  dual  valves  are 
used,  they  are  used  both  for  inlet  and  exhaust. 

In  addition  to  providing  the  maximum  free  area  through  the 
valves  it  is  important  that  the  form  of  the  passage  through  which 
the  gas  is  flowing  should  be  such  as  to  offer  minimum  resistance 
and  as  far  as  possible  to  permit  smooth  flow  without  the  forma- 


tn/et 


Exhaust- 


FIG.   115. — Trumpet-shaped  FIG.  116. — Valves  of  Basse-Selve  engine, 

valve. 

tion  of  eddies.  Especially  is  this  true  for  the  inlet  valve  as  it 
will  influence  materially  the  volumetric  efficiency  of  the  engine. 
The  limitations  inherent  in  an  engine  as  light  and  compact  as  an 
airplane  engine  will  generally  prevent  the  adoption  of  ideal 
forms  of  passage,  but  some  improvements  over  common  practice 
(as  represented  in  Fig.  113)  are  possible.  One  of  these  is  in  the 
shape  of  the  valve.  If  the  valve  is  formed  with  a  trumpet  head 
which  flows  with  a  large  radius  into  the  stem  as  in  the  Siddeley 
"Puma"  inlet  valve  (Fig.  115),  the  gas  flow  lines  will  be  consider- 
ably improved;  such  valves  should  be  hollowed  out  to  reduce  the 
weight.  The  approach  of  the  gas  to  the  valve  is  usually  smoother 
when  the  valves  are  inclined  than  when  they  are  vertical;  compare, 
for  example,  the  Basse*-Selve  (Fig.  116)  with  the  Curtiss  K  (Fig. 


158 


THE  AIRPLANE  ENGINE 


87),  both  of  which  are  good  examples  of  smooth  passages  without 
sudden  enlargements.  In  the  Liberty  engine  (Fig.  91)  the 
form  of  the  gas  passages  is  such  as  to  set  up  eddies.  On  the 
exhaust  side  the  need  for  adequate  water-jacketing  of  the  valve 
seat  and  guide  will  generally  lead  to  a  less  favorable  form  of  gas 
passage  than  on  the  inlet  side.  This  is  especially  noticeable  in  the 
Basse-Selve  engine  (Fig.  116),  in  which  the  water  jacket  is  carried 
all  round  the  whole  length  of  the  valve  stem  guide;  in  most 
engines  such  complete  jacketing  is  not  attempted. 

The  cross-section  area  of  inlet  pipes  and  ports  should  be 
approximately  the  same  as  the  valve  area  so  as  to  avoid  loss 
of  head  due  to  change  of  cross-section.  With  an  engine  using  0.6 


50 


'800        1200  1600        2000 

Revolutions  per    Minute 


2400 


1200  1600         2000        2400 

Revolutions  per  Minute 


FIG.  117. — Variation  of  engine     FIG.  118. — Variation  of  engine 
power  with  valve  lift.  power  with  valve  diameter. 

lb.  of  fuel  per  horse-power  hour  and  15  Ib.  of  air  per  pound  of  fuel, 
the  volume  of  charge  entering  the  cylinder  will  be  approximately 
120  cu.  ft.  and  of  the  exhaust  gases  250  to  300  cu.  ft.  per  horse- 
power hour.  With  a  velocity  of  150  ft.  per  second  past  the  inlet 

120 
valve  the  inlet  pipe  cross-section  is  4  X  »  ar\r\  X  144  -r-  150  = 

0.128  sq.  in.  per  horse  power.  Good  engines  show  usually  from 
0.14  to  0.16  sq.  in.  per  horse  power  for  the  inlet  pipes  and  ports. 
Exhaust  pipes  and  ports  are  usually  of  the  same  size  as  inlet 
pipes  and  ports,  but  are  occasionally  larger.  In  the  Austro-Daim- 
ler  engine  inlet  pipes  are  0.126  sq.  in.  per  horse  power  and  exhaust 
pipes  0.18  sq.  in. 

The  effect  of  valve  lift  on  engine  capacity  is  shown  in  Fig.  117, 
which  is  plotted  from  Pomeroy's1  test  results.  The  engine 
was  90  X  120  mm.  and  had  an  adjustable  inlet  valve  lift.  The 

1  The  Automobile  Engineer,  Feb.,  1919,  p.  44. 


VALVES  AND  VALVE  GEARS  159 

valve  diameter  was  1.75  in. ;  the  valve  areas  (irdh)  used  in  the  test 
were  1.8,  1.4  and  0.7  sq.  in.  As  indicated  in  the  figure  the  b.h.p. 
was  the  same  for  the  two  larger  areas  up  to  1,500  r.p.m.  above 
which  speed  the  larger  area  showed  its  superiority.  The  lowest 
lift  showed  marked  inferiority. 

Tests  of  another  engine  of  the  same  size  with  constant  lift 
and  with  the  valve  areas  kept  the  same  as  in  the  previous  engine, 
by  fitting  different  diameters  of  valves  (1%6>  1%>  and  1  in.), 
gave  the  results  shown  in  Fig.  118.  It  is  noteworthy  that  the  per- 
formances of  these  three  valves  were  practically  identical  up  to 
1,600  r.p.m.,  above  which  speed  the  smallest  valve  showed  inferior 
results  and  the  intermediate  size  valve  showed  best  results.  It 
is  seen  by  comparing  Figs.  117  and  118  that  valve  area  alone  is  not 
important;  it  is  necessary  to  know  the  lift-diameter  ratio  also. 
The  highest  curve  in  Fig.  118  is  obtained  with  a  lift  of  23.6  per 
cent  of  the  diameter;  in  Fig.  117  the  highest  curve  has  a  lift- 
diameter  ratio  of  18.7  per  cent,  but  it  is  probable  that  still  better 
results  would  have  been  obtained  with  a  higher  lift. 

Valve  Materials. — Inlet  valves  in  airplane  engines  under 
normal  operation  may  reach  temperatures  of  over  1,100°F.; 
exhaust  valves  may  go  to  1,600°F.  or  higher.  The  heat  received 
by  the  head  of  an  exhaust  valve  is  dissipated  in  three  ways: 
(1)  by  conduction  down  the  stem  to  the  guide,  (2)  by  direct 
radiation  from  the  back  surface  of  the  head  and  (3)  by  direct 
conduction  from  the  face  to  the  valve  seat.  The  last  of  these  is 
by  far  the  most  important.  To  be  effective  it  is  essential  that 
the  valve  should  have  good  metallic  contact  with  its  seat  through- 
out the  whole  of  the  explosion  stroke.  If,  through  valve  warping 
or  the  lodging  of  scale  on  the  seat,  there  should  be  any  leakage  of 
gas  past  the  valve,  there  will  be  rapid  heating  at  the  place  where 
the  leakage  occurs  and  the  valve  will  burn  away  at  that  place. 
Another  prolific  cause  of  valve  burning  is  persistent  preignitions 
in  the  cylinders;  it  is  found  that  valves  which  stand  up  satis- 
factorily under  normal  operation  fail  very  rapidly  when  per- 
sistent preignitions  occur;  with  such  preignitions  the  temperature 
of  the  exhaust  valve  may  rise  to  2,100°F.  If  the  exhaust  ports 
are  so  designed  that  the  exhaust  gases  play  directly  on  the  neck  of 
the  valve  this  may  become  highly  heated  and  may  actually 
supply  heat  to  the  valve  head  instead  of  taking  it  away;  in  such  a 
case  overheating  of  the  valve  is  likely  to  occur.  It  is  also  impor- 
tant that  the  valve  guides  should  be  efficiently  water-cooled  and 


160  THE  AIRPLANE  ENGINE 

should  not  project  into  the  exhaust  pocket  so  as  to  be  heated 
directly  by  the  exhaust  gases.  A  final  cause  of  overheating  the 
exhaust  valve  is  the  use  of  an  overrich  mixture  which  may  be 
still  burning  during  the  exhaust  stroke. 

An  interesting  suggestion  for  valve  cooling  is  the  use  of  a  hollow 
stem  into  which  is  put  a  small  amount  of  mercury  before  plugging. 
The  liquid  mercury  in  contact  with  the  hot  center  of  the  valve 
head  is  vaporized  and  is  condensed  again  in  the  upper  part 
of  the  stem.  The  mercury  thus  acts  as  a  heat  carrier  abstracting 
from  the  valve  head  its  latent  heat  each  time  it  is  vaporized. 
The  vapor  pressure  of  mercury  at  820°F.  is  50  Ib.  per  square  inch. 

The  principal  types  of  valve  failure  are  (1)  elongation  of 
the  stem,  (2)  distortion  of  the  head,  (3)  cracks  in  the  valve  face, 
(4)  wear  of  the  stem,  (5)  wear  of  the  foot,  (6)  burning  of  the 
head,  (7)  scaling  and  (8)  breaking  due  to  self -hardening.  Elon- 
gation of  the  stem  results  either  from  the  use  of  a  steel  of  insuffi- 
cient strength  at  the  working  temperature  or  from  overheating 
of  the  stem.  Distortion  of  the  head  occurs  usually  when  proper 
heat  treatment  has  not  been  given  to  the  valve  forging  before 
machining;  in  other  cases  unequal  heating  or  softening  under 
the  action  of  high  temperature  may  be  the  causes.  Cracks  come 
usually  from  cracks  in  the  steel  from  which  the  valves  are  made; 
they  are  fairly  common  and  are  dangerous  as  they  may  result  in 
the  breaking  away  of  a  section  of  the  valve.  Wear  on  the  valve 
stem  occurs  usually  in  rotary  engines  which  produce  a  side 
pressure  due  to  the  inertia  of  the  valve.  Wear  of  the  valve  foot 
results  from  the  hammering  of  the  tappet  or  the  wipe  of  the  cam; 
it  is  diminished  by  hardening  the  foot  or  by  the  use  of  a  cap. 
Burning  is  due  to  overheating. 

A  steel  to  be  satisfactory  for  exhaust  valves  in  airplane  engines 
should  have  the  following  properties  as  stated  by  Aitchison.1 

1.  The  greatest  possible  strength  at  high  temperatures. 

2.  The  highest  possible  notched  bar  value  (resistance  to  impact). 

3.  The  capacity  of  being  forged  easily. 

4.  The  capacity  of  being  manufactured  free  from  cracks. 

5.  The  capacity  of  being  easily  heat-treated. 

6.  The  least  possible  tendency  to  scale. 

7.  The  ability  to  retain  its  original  physical  properties  after  repeated 
heatings  for  prolonged  periods. 

8.  Freedom  from  liability  to  harden  on  air  cooling. 

1  The  Automobile  Engineer,  Nov.,  1920. 


VALVES  AND  VALVE  GEARS  161 

9.  Freedom  from  distorting  stresses  after  heat  treating. 

10.  Hardness  to  resist  stem  wear. 

11.  Capacity  of  being  hardened  at  the  foot. 

12.  Reasonable  ease  of  machining. 

The  best  steels  for  exhaust  valves  are  in  five  classes: 

1.  Tungsten  steel  with  not  less  than  14  per  cent  tungsten  and  about  0.6  per 
cent  carbon. 

2.  High  chromium  steels  (stainless  steel)  with  about  13  per  cent  chromium 
and  about  0.35  per  cent  carbon. 


°650 


150  850  950 

Temperature ,  Decj.  C- 


FIG.  1 19. — Resistance  of  valve  steels  to  scaling. 


3.  Steel  containing  from  7  to  12  per  cent  chromium  and  about  0.6  per  cent 
carbon. 

4.  Steels  containing  about  3  per  cent  nickel. 

5.  Ordinary  nickel-chromium  steels. 

Of  these  steels  the  first  four  are  superior  to  the  last.  The 
nickel-chromium  steels  are  difficult  to  manufacture  free  from 
flaws,  they  tend  to  harden  during  the  running  of  the  engine,  they 
scale  rapidly  and  they  show  no  superiority  at  high  temperatures 
over  the  other  steels.  The  relative  resistances  to  scaling  are 
shown  in  Fig.  119,  from  which  it  is  apparent  that  stainless  steel  is 
superior  to  the  others.  The  tensile  strengths  of  these  steels  at 

higher  temperatures  are  given  in  the  following  table. 
11 


162  THE  AIRPLANE  ENGINE 

ULTIMATE  STRENGTH  OF  VALVE  STEELS,  POUNDS  PER  SQUARE  INCH 


Steel 


Temperature,  degrees 
Fahrenheit 


1,300 


1,650 


High  tungsten,  high  carbon  

39,600 

19,700 

High  tungsten,  low  carbon 

34  700 

14,100 

High  chromium,  high  carbon  

33,800 

16,800 

High  chromium,  low  carbon 

27  100 

10,700 

Low  chromium,  high  carbon  

41,400 

16,800 

Low  chromium,  low  carbon 

38  ,  000 

15,800 

3  per  cent  nickel,  high  carbon  

25,800 

10,100 

3  per  cent  nickel,  low  carbon        .    . 

21,000 

8,700 

Nickel  chromium  

23,500 

10,100 

The  3  per  cent  nickel  steel  is  much  cheaper  than  the  others  but 
is  markedly  inferior  in  tensile  strength  at  high  temperatures  and 
consequently  should  be  used  only  on  inlet  valves  or  for  the 
exhaust  valves  of  rotary  engines.  The  high  chromium  (stainless) 
steel  is  highly  resistant  to  scaling  and,  if  of  low  carbon  content, 
is  readily  machined  but  is  not  easy  to  forge  and  is  liable  to 
cracks.  High  tungsten  steel  retains  its  strength  best  of  any  steel 
at  high  temperatures  and  is  fairly  resistant  to  scaling.  Exhaust 
valves  which  are  liable  to  be  subjected  to  unusually  high  tempera- 
tures should  be  of  tungsten  steel;  for  more  moderate  tempera- 
tures stainless  steel  will  be  more  durable.  Monel-metal  valves 
have  been  used,  and  although  they  have  stood  up  well  under  test 
on  the  Hispano-Suiza  engine,  they  have  failed  rapidly  on  the 
Liberty  engine. 

Valve  Operation. — The  valves  of  modern  airplane  engines  are 
mechanically  operated;  automatic  action,  which  is  found  in  some 
automobile  engines  and  in  a  few  of  the  earlier  airplane  engines, 
must  always  result  in  lowered  volumetric  efficiency  and  capacity. 
Actuation  of  the  valves  is  by  means  of  cams  acting  either  directly 
or  indirectly.  The  camshafts  may  be  placed  near  the  base  of  the 
cylinders  and  operate  the  valves  through  push  rods  and  rocker 
arms,  or  overhead  camshafts  may  be  used  acting  on  the  valves 
directly  or  through  rocker  arms. 


VALVES  AND  VALVE  GEARS 


163 


A  good  example  of  push-rod  operation  is  shown  in  the  Benz 
230  engine  (Fig.  78),  which  has  separate  camshafts  for  the  inlet 
and  exhaust  valves,  both  located  in  the  crankcase;  a  similar 
arrangement  is  used  in  the  May  bach  engine  (Fig.  81).  In  Vee 
engines  the  usual  practice,  where  push  rods  are  employed,  is  to 
have  a  single  camshaft,  located  in  the  angle  of  the  Vee,  inside  the 
crankcase,  carrying  inlet  and  exhaust  cams  for  both  rows  of 
cylinders;  the  Curtiss  OX  and  V2  engines  (Figs.  60  and  62) 
show  this  arrangement,  which  is  also  used  on  the  Benz  300  (Fig. 
131).  In  recent  years  the  tendency  has  been  to  do  away  with 
push  rods  and  to  use  overhead  camshafts.  This  last  arrangement 
reduces  the  weight  and  complexity  of  the  valve  gear,  and,  in 
consequence  of  the  smaller  number  of  joints  involved,  makes  for 
better  maintenance  of  the  valve  timing.  There  may  be  either 
(1)  one  camshaft  over  each  row  of  cylinders  acting  directly  on 
the  valves  as  in  the  Hispano-Suiza  (Fig.  51),  or  (2)  one  camshaft 
acting  directly  on  one  set  of  valves  and  indirectly  through  a 
rocker  arm  on  the  other  set  as  in  the  Siddeley  "Puma"  (Fig.  8),  or 
(3)  one  camshaft  acting  indirectly  through  rocker  arms  on  both 
sets  of  valves  as  in  the  Liberty  (Fig.  47),  Basse-Selve  (Fig.  129), 
Bugatti  (Fig.  67),  Fiat  (Fig.  76),  Rolls-Royce  (Fig.  70),  Mercedes, 
Lorraine-Dietrich,  Renault,  Austro-Daimler,  or  (4)  two  camshafts 
acting  directly  on  the  two  sets  of  valves  as  in  the  Curtiss  K 
(Fig.  56)  and  Napier  "Lion"  (Fig.  73). 

Cams. — The  shape  of  the  cam  depends  on  the  desired  valve 
movement  and  on  the  form  and  location  of  the  cam  follower. 


FIG.  120. — Individual  cam. 


(a)  (b)  (c) 

FIG.  121. — Cam  forms. 


The  cams  are  usually  integral  with  the  camshaft,  which  gives 
maximum  security  and  accuracy  of  location,  but  sometimes  they 
are  fastened  by  taper  pins  to  the  hollow  camshaft,  as  in  Fig.  120; 
this  arrangement  permits  more  satisfactory  hardening  of  the  cams 
and  the  replacement  of  a  worn  cam  but  is  less  secure  and  may 
become  slack.  The  cams  are  sometimes  made  with  convex 
flanks  as  in  Fig.  1216,  or  with  flat  surfaces  tangential  to  circular 


164 


THE  AIRPLANE  ENGINE 


arcs  as  in  Fig.  121a,  or  with  flanks  that  change  from  concave  to 
convex  and  a  top  which  is  concentric  with  the  camshaft  as  in 
the  constant  acceleration  cam  of  Fig.  121c.  The  cam  follower 
may  be  flat  as  in  Fig.  122a,  rounded  as  in  Fig.  1226,  or  a  roller 
as  in  Fig.  122c,  and  it  may  be  fixed  on  the  end  of  the  valve 
plunger  or  it  may  be  mounted  on  a  radius  rod  as  in  Fig.  122cL 
With  a  flat  follower  a  convex  flanked  cam  is  used;  tangential 
and  constant  acceleration  cams  are  used  with  the  other  types  of 
follower  shown  in  Fig.  122. 

The  work  which  a  cam  has  to  do  is  in  three  parts.  (1)  It 
must  overcome  the  difference  of  gas  pressures  on  the  two  sides  of 
the  valve.  This  pressure  difference  is  important  only  in  the  case 
of  the  exhaust  valve,  which  just  previous  to  opening  has  a  pressure 


FIG.  122. — Cam  followers. 

of  about  60  Ib.  per  square  inch  (gage)  on  one  side  and  atmospheric 
pressure  on  the  other.  With  a  valve  2%  in.  diameter  the  pres- 
sure difference  is  nearly  300  Ib.  at  the  instant  of  opening  and  falls 
rapidly  to  a  negligible  quantity.  (2)  The  valve  spring  is  operat- 
ing at  all  times  to  keep  the  valve  closed;  the  compression  on 
the  spring  is  usually  about  50  Ib.  when  the  valve  is  closed  (see 
table  4) .  The  cam  must  do  work  in  compressing  the  spring.  (3) 
The  moving  parts  from  the  cam  to  the  valve,  including  the  valve 
and  spring,  must  be  accelerated  and  work  must  be  done  in  giving 
them  the  necessary  acceleration.  The  force  required  to  acceler- 
ate these  parts  is  determined  by  the  design  of  cam  and  follower 
and  by  the  masses  that  have  to  be  accelerated. 

For  maximum  volumetric  efficiency  of  the  engine  the  valves 
should  open  promptly,  should  remain  wide  open  as  long  as 
possible  and  then  should  close  promptly.  If  a  valve  is  to  be 
opened  in  a  given  time  (number  of  degrees  of  crankshaft  rotation) 


VALVES  AND  VALVE  GEARS  165 

the  force  required  to  accelerate  the  moving  parts  will  be  kept  a 
minimum  by  making  the  acceleration  constant  and  thereby 
keeping  the  accelerating  force  constant.  During  the  opening 
the  moving  parts  must  be  first  accelerated  and  then  brought  to 
rest;  the  deceleration  is  accomplished  by  the  valve  spring  and, 
sometimes,  if  a  push  rod  is  used,  by  an  additional  spring  acting 
on  the  push  rod.  Smooth  action  will  be  obtained  when  the 
deceleration  is  constant  and  has  the  same  value  as  the  accelera- 
tion. The  cam  does  not  necessarily  do  any  work  at  all  during 
the  decelerating  period. 

The  acceleration  and  the  force  required  to  produce  it  are 
readily  calculable.  Suppose  the  valve  to  move  from  the  closed 
to  the  fully  open  position  in  60  deg.  of  crankshaft  rotation  and 
that  the  moving  parts  are  accelerating  uniformly  for  half  that 
period,  or  30  deg.;  and  are  decelerating  uniformly  for  the  next 
30  deg.  Then  at  1,500  r.p.m.  the  time,  t,  available  for  acquir- 

60          30         1 
ing  maximum  velocity  is  -        X  TT    =         sec.     If  the  lift  is 


0.5  in.  the  distance,  d,  moved  in  this  time  is  0.25  in,  and  the 
acceleration,  a,  is  given  by  d  =  ~at2,  or  a  =  3,750  ft.  per  second 

A 

per  second.  If  the  weight,  w,  of  the  moving  parts  is  1  Ib.  and 
if  all  of  the  parts  move  with  the  same  velocity  as  the  valve, 

the  force  required  to  accelerate  the  parts  will  be  —  a  =  110  Ib. 

y 
The  force  exerted  by  the  cam  will  be  greater  than  this  by  an 

amount  equal  to  the  valve  spring  compression  and,  at  the  instant 
of  opening  the  exhaust  valve,  by  the  gas-pressure  difference. 
With  the  numerical  values  given  above  the  force  exerted  by  the 
exhaust  cam  at  the  moment  when  the  valve  begins  to  open  must 
be  110  +  300  +  50  =  460  Ib.  As  there  is  always  some  tappet 
clearance  to  permit  expansion  of  the  valve  stem  without  forcing 
the  valve  to  lift,  this  maximum  force  occurs  a  short  time  after 
the  cam  has  come  into  action.  The  force  will  diminish  rapidly 
as  the  gas  pressure  in  the  cylinder  falls  but  will  tend  to  increase 
later  with  increasing  spring  compression.  During  the  decelerat- 
ing period  the  spring  pressure  would  have  to  be  greater  than  110 
Ib.  to  bring  the  valve  to  rest  in  30  deg.  of  crank  rotation.  During 
the  closing  of  the  valve  the  gas  pressure  difference  is  absent,  the 
acceleration  is  due  to  spring  action  and  the  deceleration  is 
brought  about  by  pressure  on  the  cam. 


166 


THE  AIRPLANE  ENGINE 


The  forces  exerted  by  the  cam,  which  have  just  been  considered, 
are  the  radial  forces,  R,  acting  along  the  push  rod,  or,  in  the  case 
of  overhead  camshafts,  at  right  angles  to  the  outer  end  of  the 
rocker  arm.  The  actual  pressure,  N,  between  the  cam  and  its 
follower  acts  normal  to  the  surface  of  contact  and  will  be  greater 
than  the  radial  force  throughout  the  accelerating  period.  The 
relation  between  these  two  forces  is  indicated  by  the  triangle  of 
forces  in  Fig.  1226.  The  side  thrust,  S,  may  be  trouble- 
some. The  quicker  the  opening  of  the  valve  the  greater  will  be 
the  acceleration  force,  R,  and  the  greater  will  be  the  ratio  of  both 
N  and  S  to  R.  In  other  words,  the  normal  pressure  and  the 
side  thrust  increase  much  more  rapidly  than  the  radial  force. 
The  side  thrust  is  particularly  objectionable  with  valve  plunger 
guides  as  in  Figs.  1226  and  c;  with  the  arrange- 
ment of  Fig.  I22d,  or  with  the  cam  operating 
directly  on  the  rocker  lever,  a  rapid  operjing  of 
the  valve  can  be  obtained  without  trouble 
from  side  thrust. 

A  good  example  of  the  constant  acceleration 
type  of  cam  with  roller  follower  is  shown  in  the 
Maybach  engine,  Fig.  123.  The  displacement, 
velocity,  and  acceleration  curves  both  for  inlet 
opening  and  for  exhaust  closure  are  given  in 
Fig.  124;  it  will  be  seen  that  the  valves  open 
and  close  rapidly,  and  remain  full  open  for 
considerable  periods  of  time.  The  velocity 
of  the  exhaust  valve  when  closing  increases 
uniformly  for  60  deg.  of  crank  rotation,  then 
decreases  but  not  quite  uniformly  for  the 
next  46  deg.;  the  inlet  valve  on  opening  has 
acceleration  which  increases  for  about  48  deg., 
when  maximum  velocity  is  attained,  and  then 
comes  to  rest  after  60  deg.  more  of  uniform 
deceleration.  Valve  displacements  for  the 
whole  cycle  are  shown  in  Fig.  125;  it  will  be  seen  that  the  valves 
are  wide  open  for  considerable  fractions  of  the  stroke. 

The  tangential  cam  is  used  in  the  Liberty  engine  (Fig.  130) 
operating  on  a  roller  at  the  end  of  the  rocker  arm.  With  this 
type  of  cam,  the  center  of  curvature  of  the  highest  part  of  the 
cam  cannot  coincide  with  the  center  of  the  camshaft  (as  in  the 
constant  acceleration  cam),  and  consequently  the  valve  cannot 


FIG.    123.  —  May- 
bach  valve  gear. 


VALVES  AND  VALVE  GEARS 


167 


stay  at  its  wide-open  position.  The  actual  valve  lifts  are  shown 
in  Fig.  126  plotted  against  crank  position,  and  in  Fig.  127 
plotted  against  piston  position.  The  valve  opening  is  not  so 
good  as  in  the  Maybach  engine,  but  the  forces  required  to  acceler- 

Exhaust  DISPLACEMENT      CURVE  Inlet 

Crankshaft    Degrees 
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FIG.   124. — Displacement,  velocity  and  acceleration  curves  for  the  valves  of  the 

Maybach  engine. 

ate  the  moving  parts  are  less  in  consequence  of  the  longer  time 
available  for  opening  or  closing  the  valve;  with  symmetrical 
cams  this  time  is  one-half  the  total  time  the  valve  is  open.  In 
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Diagram  of  Exhaust  and  Inlet  Port  Opening  in  Relation  to  Piston  Position 

FIG.  125. — Valve  openings  of  Maybach  engine. 

29  deg.;  in  the  Liberty  engine  the  inlet  valve  is  open  for  215  deg. 
and  the  valve,  while  opening,  is  accelerating  for  54  deg.  of  crank 
rotation.  As  the  acceleration  is  inversely  as  the  square  of  the 
time  taken  to  lift  the  valve  through  a  given  distance,  the  force 


168 


THE  AIRPLANE  ENGINE 


required  to  overcome  the  inertia  of  the  moving  valve  parts  would 
be  3.5  times  as  great  for  an  engine  using  29  deg.  of  crank  rotation 
for  the  valve  acceleration  as  for  the  same  engine,  with  the  same 
revolutions  per  minute,  using  54  deg.  of  crank  rotation. 

Valve  Springs. — The  function  of  the  valve  spring  is  to  deceler- 
ate the  valve  moving  parts  during  the  latter  half  of  the  valve 
opening  and  to  accelerate  them  during  the  first  half  of  the  valve 


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FIG.  126. — Valve  lifts  of  Liberty  engine  plotted  against  crank  positions. 

closure.  In  addition,  the  exhaust  valve  spring  must  be  strong 
enough  to  keep  the  valve  closed  during  the  suction  stroke  when 
the  engine  is  idling.  During  idling  the  pressure  in  the  cylinder 
may  be  10  Ib.  per  square  inch  below  atmospheric  pressure,  which, 
with  a  2%-m.  valve,  gives  a  pressure  difference  of  50  Ib.  between 
the  front  and  back  of  the  valve.  The  spring  pressure  required 
for  acceleration  is  normally  in  excess  of  that  required  to  keep  the 
exhaust  closed  during  idling. 


VALVES  AND  VALVE  GEARS 


169 


Cylindrical  helical  springs  are  employed  almost  universally, 
but  occasionally  conical  helical  springs  may  be  used  (Fig.  133),  or, 
when  minimum  height  is  required,  the  rat-trap  type  of  spring,  as 
in  the  Curtiss  engine  (Fig.  128).  In  the  larger  engines  two  con- 
centric cylindrical  helical  springs  are  used  as  in  the  Liberty  engine 


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Diagram  of  Exhaust  and   Inlet  Port  Opening  in  Return  to  Piston  Position 

FIG.  127. — Valve  lifts  of  Liberty  engine  plotted  against  piston  positions. 

(Fig.  130)  and  the  Benz  300  inlet  (Fig.  131),  or  even  three 
springs  as  in  the  Bugatti  exhaust  valve  (Fig.  67).  An  advan- 
tage of  multiple  springs  is  that,  in  case  of  breakage  of  one  spring, 
the  valve  cannot  fall  into  the  cylinder. 


FIG.   128. — "Rat  trap"  valve  spring  (Curtiss  engine). 

The  maximum  safe  working  load,  P,  in  pounds,  on  a  cylindrical 
helical  spring  of  outside  diameter  D  in.  and  with  steel  of  diameter 
d  in.  is  given  by 


170 


THE  AIRPLANE  ENGINE 


where  S  is  the  safe  shearing  stress  of  the  material  in  pounds 
per  square  inch.  The  value  of  S  varies  from  about  80,000  to 
150,000  (increasing  as  the  diameter  of  the  wire  diminishes)  for 
springs  that  are  used  intermittently.  For  continuous  use,  as  in 
an  airplane  engine,  about  half  these  values  should  be  employed, 
or,  for  the  usual  sizes  of  spring  steel,  about  60,000  Ib. 

The  deflection,  /,  in  inches  of  one  coil  of  a  cylindrical  helical 
spring  is  given  by 

SnPD* 


where  G  is  the  modulus  of  elasticity  in  shear  and  may  be  taken  as 
12,000,000  Ib.  per  square  inch.  The  following  table  gives  safe 
loads,  Pt  and  the  corresponding  deflection,  /,  of  one  coil  for  various 
springs.  It  is  calculated  for  S  =  60,000  Ib.  per  square  inch. 

SAFE  LOADS  AND  DEFLECTION  OF  CYLINDRICAL  HELICAL  SPRINGS 


Outside  diameter  of  spring,  D  in. 


YVIIC  gage 

1.75 

2.00 

2.25 

2.50 

No.  8  

0  162 

P 

64. 

55  5 

48  8 

43  5 

No.  6  

0  192 

f 
P 

0.23 
107. 

0.33 
92.5 

0.42 

81 

0.57 

72 

No.  5  

0  205 

/ 
P 

0.20 
131. 

0.26 
113. 

0.34 
99. 

0.42 

88  5 

No.  4  

0  225 

f 
P 

0.18 
175. 

0.25 
150. 

0.32 
132. 

0.36 
118. 

No.  3  

0  242 

f 
P 

0.16 
225. 

0.22 
195. 

0.29 
.170. 

0.36 
152. 

f 

0.11 

0.16 

0.20 

0.30 

For  square  steel  of  side  d  in.  the  tabular  valves  of  P  should  be 
multiplied  by  1.2,  and  the  values  of/  by  0.59. 

The  valve  spring  retainer  is  usually  a  washer  or  cupped  disc 
with  a  downwardly  turned  flange  to  center  the  spring.  The 
retainer  is  fastened  to  the  valve  stem  in  various  ways.  It  is 
sometimes  held  by  a  nut  which  screws  on  to  the  threaded  upper 
end  of  the  valve  stem  and  is  locked  by  a  split  pin  as  in  the  Basse- 
Selve  engine  (Fig.  129);  or  it  is  held  by  a  cotter  through  the 
valve  stem,  locked  in  position  by  wire  clips  as  in  the  Maybach 
engine  (Fig.  123);  or  the  valve  stem  is  turned  to  a  smaller  diame- 
ter for  a  short  length  near  the  top  and  held  by  a  conical  split 


VALVES  AND  VALVE  GEARS 


171 


collar  as  in  the  Liberty  engine  (Fig.  130),  the  Benz  300  (Fig. 
131)  and  the  Fiat  engine  (Fig.  134).  In  those  engines  in  which 
the  cam  acts  directly  on  a  flat  cam  follower  on  top  of  the  valve 
stem  this  follower  may  serve  to  hold  the  spring  retainer.  In 
the  Hispano-Suiza  engine  (Fig.  132)  the  upper  end  of  the  valve 
stem  is  slotted  to  receive  the  disc-shaped  retainer  and  is  threaded 
internally.  The  retainer  slips  over  the  valve  stem  and  is  pro- 
vided with  a  key  which  fits  into  the  slot  and  prevents  rotation; 
it  has  fine  notches  on  its  upper  surface  to  mesh  with  correspond- 
ing notches  on  the  lower  face  of  the  follower.  The  follower  stem 
screws  into  the  hollow  valve  stem  and  after  being  screwed  into 
the  desired  position  the  retainer  is  permitted  to  come  into  contact 


FIG.  129. — Valve  gear  of  Basse-Selve  engine. 

with  it  and  locks  it  in  position.  A  similar  arrangement  is  shown 
for  the  Siddeley  "Puma"  engine  (Fig.  133),  but  in  this  case  a 
volute  steel  spring  of  rectangular  section  is  used. 

With  dual  valves  one  spring,  or  a  pair  of  concentric  springs, 
can  be  made  to  serve  for  the  two  valves  as  in  the  Fiat  engine 
(Fig.  134).  The  springs  are  in  a  steel  yoke  or  cage.  The  inner 
spring  is  mounted  on  a  central  guide  tube  while  the  outer  spring 
is  retained  by  the  yoke.  The  lugs  on  the  two  sides  of  the  yoke 
fit  over  split  locking  cones  which  are  held  by  a  spring  ring  against 
the  grooves  turned  in  the  upper  ends  of  the  valve  stems. 

Rocker  arms  are  usually  pivoted  in  plain  bearings  except 
in  German  practice,  where  ball  or  roller  bearings  are  commonly 


172 


THE  AIRPLANE  ENGINE 


used,  as  in  the  Basse-Selve  engine  (Fig.  129)  and  the  Benz  300 
(Fig.  131).  In  this  last 
engine,  which  has  one  in- 
take and  two  exhaust 
valves,  the  rocker  for  the 
intake  valve  is  brought 
obliquely  under  the 
double  rocker  of  the  ex- 
haust valves  and  the  push 
rods  are  thereby  kept  in 
line  and  close  to  the  cyl- 
inders without  unduly 
shortening  the  length  of 
the  rocker.  Double 
springs  of  very  low  height 
are  required  for  the  in- 
take valve  and  are  sunk 
in  the  dished  cylinder 
head. 


FIG.  130.— Valve  of  Liberty 
engine. 


FIG.  131. — Valve  gear  of  Benz  300. 


Adjustment  has  to  be  provided  for  the  tappet  clearance.     A 
common  method  is  that  shown  for  the  Hall  Scott  ATA  engine 


VALVES  AND  VALVE  GEARS 


173 


(Fig.  135),  in  which  the  hardened  steel  set  screw  at  the  end  of  the 
rocker  arm  is  clamped  in  position  by  a  lock  screw  at  the  end  of 
the  split  rocker  arm;  in  this  engine  the  clearance  is  0.02  in.  when 
the  engine  is  cold.  In  the  Liberty  engine  (Fig.  130)  the  adjusting 


FIG.  132. — Valves  of  Hispano-Suiza  engine. 

screw  is  held  in  place  by  a  lock  nut.  The  amount  of  clearance 
should  exceed  the  expansion  of  the  valve  stem  and  depends 
mainly  on  the  length  of  the  valve  stem.  As  the  inlet  valve  is 
colder  than  the  exhaust  valve  it  requires  less  tappet  clearance; 
in  the  Liberty  engine  the  inlet  clearance  is  0.015  in.,  the  exhaust 


FIG.  133. — Valves  of  Siddeley  "Puma"  engine. 

0.020  in.     Too  much  clearance  is  to  be  avoided,  as  producing 
noise  and  possible  breakage  of  parts. 

Valve  Timing. — If  the  inlet  and  exhaust  valve  openings  could 
be  made  unobstructed  and  large  enough,  and  if  the  openings 
and  closings  were  sufficiently  rapid,  each  valve  could  be  timed  to 


174 


THE  AIRPLANE  ENGINE 


open  at  the  beginning  of  a  piston  stroke  and  to  close  at  the  end 
of  that  stroke.  In  actual  engines  it  is  necessary  for  the  valves  to 
depart  from  this  timing  in  the  interests  of  high  volumetric  effi- 
ciency and  capacity.  Especially  must  the  exhaust  valve  open 
early  and  the  inlet  valve  close  late.  The  time  of  opening  of  the 


FIG.  134. — Valves  of  Fiat  engine. 


inlet  valve  is  related  to  the  time  of  closing  of  the  exhaust  valve : 
the  exhaust  valve  usually  closes  completely  shortly  after  the 
inlet  valve  starts  to  lift.  The  valve  does  not  start  to  open  until 
the  tappet  has  moved  a  distance  equal  to  the  tappet  clearance. 
The  inlet  valve  opening  is  usually  10  to  15  deg.  past  top  dead 
center  (T.D.C.)  but  sometimes  occurs  before  top  dead  center. 


Fia.  135. — Tappet  adjustment,  Hall-Scott  engine. 

Inlet  valve  closure  is  usually  about  40  deg.  past  bottom  dead 
center  (B.D.C.),  which  corresponds  to  about  10  per  cent  of  the 
return  stroke  of  the  piston  and  causes  a  corresponding  decrease 
in  the  compression  ratio.  The  exhaust  valve  opens  from  45  to 
50  deg.  before  bottom  dead  center  and  closes  about  10  deg.  past 
top  dead  center.  The  actual  timing  for  any  engine  should  be 


VALVES  AND  VALVE  GEARS 


175 


determined  by  operation  of  the  engine  and  observation  of  the 
timing  giving  maximum  capacity.  The  inlet  valve  should  not  be 
opened  until  a  partial  vacuum  is  established  in  the  cylinder;  too 
late  a  closure  will  result  in  reduction  of  compression  below  that 
obtainable  with  proper  timing.  Valve  timings  for  various 
engines  are  given  in  the  following  table: 

EXAMPLES  OF  VALVE  TIMING 


Engine 

Inlet 

Exhaust 

Duration  of  opening 

Opens 

Closes 

Opens 

Closes 

Intake 

Exhaust 

Hall  Scott—  A7a  
Hall  Scott—  A5a  
Curtiss—  90  
Liberty  .  . 

15°L 
15°L 
12°L 
10°L 
TDC 
TDC 
10°L 
10°E 
TDC 
2°L 
8°E 
5°L 
10°-15°L 

40°L 
45°L 
40°L 
45°L 
37°L 
45°L 
52M°L 
62°L 
40°L 
51°L 
35°L 
45°L 
35°-50°L 

45°E 
50°E 
45°E 
50°E 
47°E 
47H°E 
48°E 
62M°E 
40°E 
52°E 
33°E 
55°E 
45°E 

10°L 
10°L 
TDC 
10°L 
TDC 
17H°L 
10°L 
29K°L 
10°L 
16H°L 
7°L 
18°L 
10°L 

205° 
210° 
208° 
215° 
217° 
225° 
222H° 
243° 
220° 
229° 
223° 
230° 
200°-220° 

235° 

240° 
225° 
240° 
227° 
245° 
238° 
272° 
230° 
248K° 
220° 
253° 
235° 

Curtiss  —  K12 

King-Bugatti  
Hispano-Suiza  —  180.  .  .  . 
Hispano-Suiza  —  300.  .  .  . 
Mercedes  —  '180  
Mercedes  —  240  
Maybach  —  300  

Benz  —  200  

Average  

E  =  early.     L  =  late.     TDC  =  top  dead  center. 


CHAPTER  VIII 
RADIAL  AND  ROTARY  ENGINES 

Radial  and  rotary  engines  are  characterized  by  having  the 
cylinders  disposed  at  equal  angular  intervals  around  a  complete 
circle.  The  number  of  cylinders  may  range  from  3  to  20  or 
more.  It  is  not  possible  to  arrange  more  than  10  or  11  cylinders 
in  a  circle  without  increasing  the  size  of  the  crankcase  to  dimen- 
sions which  give  an  over-all  diameter  too  large  for  use  in  an 
airplane.  If  more  cylinders  are  desired  they  have  to  be  arranged 
in  two  planes  or  banks,  with  an  equal  number  of  cylinders  in 
each  plane;  with  air  cooling,  the  cylinders  of  the  rear  ro*w  are 
staggered  with  reference  to  those  of  the  front  row;  with  water 
cooling,  they  may  lie  exactly  behind  those  of  the  front  row. 

Fixed-radial  engines  have  stationary  cylinders  and  a  revolving 
crankshaft;  there  are  usually  as  many  cranks  as  there  are  rows  of 
cylinders,  although  two  cranks  have  been  used  with  a  single  row 
of  cylinders. 

Rotary  engines  have  rotating  cylinders  and  a  fixed  crank- 
shaft; in  this  case  the  propellor  hub  is  attached  to  the  rotating 
crankcase. 

Double-rotary  engines  have  both  cylinders  and  crankshaft 
rotating  but  in  opposite  directions.  With  this  type,  two  arrange- 
ments are  possible  for  utilizing  the  power  developed:  (a)  two 
propellers  may  be  used  mounted  on  the  crankcase  and  the  crank- 
shaft respectively  and  therefore  rotating  in  opposite  directions 
but  with  right-  and  left-hand  pitches  respectively,  so  that  both 
give  thrust  in  the  same  direction;  (6)  the  crankcase  may  be  geared 
to  the  crankshaft  (with  reversal  of  motion)  and  the  power 
absorbed  by  a  single  propeller  on  the  crankshaft  or  crankcase. 

Air  Cooling. — The  disposition  of  the  cylinders  of  a  radial  or 
rotary  engine,  in  a  plane  at  right  angles  to  the  wind  and  the  slip 
stream,  gives  these  types  a  unique  opportunity  for  direct  air 
cooling  of  the  cylinders.  This  is  especially  the  case  with  rotary 
engines,  which  churn  through  the  air  as  well  as  meeting  the 
incoming  wind. 

With  cylinders  in  line,  as  in  multicylinder  vertical  or  Vee 
engines,  air  cooling  is  practicable  only  with  a  blower  for  supplying 

176 


RADIAL  AND  ROTARY  ENGINES  177 

the  cooling  air  and  with  a  system  of  ducts  for  distributing  the 
air  to  the  different  cylinders.  Renault  Freres  have  built  engines 
of  this  type  up  to  12-cylinder  Vees  but  there  are  considerable 
difficulties  in  obtaining  high  mean  effective  pressures  and  low  fuel 
consumption  with  such  cooling  as  can  be  obtained  in  this  manner. 
This  type  has  not  met  with  general  favor. 

Radial  engines  are  often  water-cooled  although  this  practice 
sacrifices  one  of  the  most  important  potential  advantages  of 
the  engine.  Rotary  engines  are  always  air-cooled  and  could  not 
be  readily  water-cooled,  both  as  a  result  of  the  mechanical  diffi- 
culties in  getting  water  to  and  from  the  rotating  cylinders,  and 
also  because  of  the  excessive  centrifugal  stresses  which  would  be 
set  up  in  the  connections  between  the  cylinder  and  the  crankcase 
as  a  result  of  the  increased  mass  of  the  cylinder  with  its  water  and 
jacket. 

Advantages  of  Radial  and  Rotary  Engines. — The  primary 
advantage  of  radial  and  rotary  engines  over  other  types  is  in 
small  weight  per  horse  power.  This  results  from  two  principal 
causes. 

1.  The  engine  can  be  air-cooled,  and  the  water  pump,  water, 
jackets,  water  pipes  and  radiator  eliminated.     The  only  impor- 
tant additional  weight  is  that  of  the  cooling  fins  on  the  cylinder. 

2.  The   crankcase  and  crankshaft  are  much  shortened  and 
lightened;  the  big  ends  of  the  connecting  rods  are  also  lighter. 

The  net  result  of  these  reductions  is  a  decrease  of  almost  40  per 
cent  in  weight  per  horse  power  of  the  power  plant.  The  Liberty 
engine  with  water  and  radiator  weighs  about  2.6  Ib.  per  horse 
power;  large  air-cooled  fixed  radial  engines  have  a  weight  of 
about  1.6  Ib.  per  horse  power. 

Other  advantages  of  the  radial  and  rotary  engines  are  short 
over-all  length,  which  permits  better  location  of  gasoline  tanks, 
pilots,  etc.;  immediate  accessibility  of  the  cylinder  heads,  and 
easy  accessibility  of  the  rest  of  the  engine  on  removal  of  the 
cowling;  ease  of  mounting  on  and  detaching  from  the  fuselage, 
the  attachment  being  to  a  vertical  plate ;  lowered  air  resistance  in 
small  sizes  which  can  be  accommodated  without  special  enlarge- 
ment of  the  fuselage — in  larger  powers  it  is  necessary  to  increase 
the  size  of  the  front  of  the  fuselage  and  thereby  increase  its  resis- 
tance until  it  is  likely  to  be  as  great  as  that  of  a  water-cooled 
engine  and  its  radiator;  engine  balance  as  good  as  in  the  best 

arrangements  of  vertical  and  Vee  engines. 
12 


178  THE  AIRPLANE  ENGINE 

Disadvantages. — The  principal  disadvantage  of  the  air- 
cooled  radial  and  rotary  engines  up  to  the  present  time  has  been 
a  lower  m.e.p.  and  higher  fuel  consumption  than  in  water- 
cooled  engines.  These  features  are  especially  true  of  the  rotary 
engine,  which  suffers  the  further  disability  that  its  revolutions  per 
minute  must  be  low  in  order  to  keep  down  the  air-churning 
resistance  and  the  centrifugal  forces  exerted  by  the  cylinders 
on  the  crankcase.  With  improved  constructions  of  cylinders 
(see  p.  202)  it  is  probable  that  the  performance  of  fixed  air- 
cooled  radial  engines  will  not  be  markedly  different  from  that 
of  water-cooled  engines.  Other  disadvantages  are  larger  oil 
consumption,  which  is  particularly  marked  in  rotary  types  but 
can  be  kept  down  in  fixed  radials;  large  over-all  diameter  neces- 
sary for  large  powers,  requiring  an  enlarged  fuselage  and  possibly 
limiting  the  power  developable  per  unit ;  large  crankpin  pressures 
in  fixed  radials;  large  gyroscopic  effect  in  rotaries,  affecting  the 
maneuvering  qualities  of  the  plane;  large  inertia  effect  in  rotaries, 
retarding  the  speeding  up  of  the  engine  on  opening  the  throttle 
but  giving  more  uniform  speed  at  low  speeds  or  with  missing 
cylinders. 

Number  of  Cylinders  and  Firing  Order. — With  a  single  crank, 
the  number  of  cylinders  of  a  radial  or  rotary  engine  must  be 
odd;  the  firing  order  will  follow  the  direction  of  rotation  of  the 
crank  in  fixed  radials  and  will  be  in  the  opposite  direction  to 
rotation  of  the  cylinders  in  rotaries.  In  both  cases  it  will  skip 
alternate  cylinders  and  will  have  occurred  in  all  the  cylinders  in 
two  complete  revolutions  of  the  crank  or  cylinders.  If  the  cylin- 
ders are  numbered  serially  the  firing  order  will  be  1,  3,  5,  7  .  .  . 
2,  4,  6,  8  ... 

With  a  two-throw  crank  and  equal  angular  spacing  of  the 
cylinders  the  number  of  cylinders  acting  on  each  crank  should  be 
odd;  the  total  number  of  cylinders  will  be  even.  (The  Smith  10- 
cylinder  engine1  with  one  row  of  cylinders  has  a  two-throw  crank 
so  that  it  is  really  equivalent  to  two  rows  of  five  cylinders  each.) 
The  firing  should  occur  alternately  on  the  two  cranks  which  are 
at  180  deg.  The  firing  order  for  regular  impulses  for  a  10- 
cylinder  engine,  with  serial  numbering,  will  be  1,  8,  5,  2,  9,  6,  3,  10, 
7,  4.  The  20-cylinder  Anzani  air-cooled  rotary  has  four  rows  of 
five  cylinders  each.  This  arrangement  keeps  the  over-all  diameter 
small  but  impairs  the  cooling  of  the  two  rear  rows. 

1  Jour.  S.  A.  E.,  Jan.,  1919. 


RADIAL  AND  ROTARY  ENGINES 


179 


With  a  two-throw  crank  and  equal  spacing  of  cylinders  in 
each  row  but  with  the  cylinders  of  the  second  row  immediately 
behind  those  of  the  first  row1  the  number  of  cylinders  in  each  row 
may  be  odd  or  even.  Firing  is  alternately  from  front  to  back  row 
cylinders  except  when  there  is  an  even  number  of  cylinders  in 
each  row.  In  that  case  two  successive  impulses  will  come  on  one 
crank  at  the  end  of  each  revolution  but  the  angle  between 
impulses  will  be  constant. 

Rotary  Engines. — In  a  rotary  engine  (Fig.  136)  the  cylinders 
rotate  about  the  crankshaft  as  a  center;  the  pistons  rotate  about 


5  4 

FIG.  136. — Diagram  of  rotary  engine. 

the  crankpin  as  a  center.  The  angular  velocity  of  the  pistons 
about  the  crankshaft  as  center  is  constant  since  it  must  be  the 
same  as  that  of  the  cylinders  which  are  rotating  at  uniform  veloc- 
ity; the  angular  velocity  of  the  pistons  about  the  crankpin  as 
center  would  be  constant  only  when  the  connecting  rods  were 
infinitely  long  or  the  crank  throw  mfinitesimally  small.  The 
result  of  the  rotations  about  the  two  centers  is  to  give  the  piston 
a  reciprocating  motion  relative  to  the  cylinders  as  shown  in 
Fig.  136  (an  analysis  of  this  motion  is  given  on  p.  507).  As  the 
cylinders  rotate  the  pistons  will  assume  the  positions  shown  rela- 
tive to  the  cylinders.  If  the  crank  is  fixed  with  its  throw  verti- 
cally upward  the  piston  will  always  be  at  the  in  dead  center 
1  See  20-cylinder  water-cooled  Anzam,  Aviation,  Feb.  15,  1920. 


180  THE  AIRPLANE  ENGINE 

when  the  cylinder  reaches  the  position  1;  it  will  be  at  the  out 
dead  center  for  the  position  vertically  below  1.  One  revolution 
of  the  cylinders  completes  one  in  and  out  stroke  of  each  piston. 

Certain  special  problems  arise  in  the  construction  of  the  rotary 
engine.  One  of  the  most  important  of  these  is  the  accommoda- 
tion of  seven  or  nine  connecting  rods  on  one  crankpin;  this 
problem  occurs  also  with  fixed-radial  engines.  Another  is  the 
attachment  of  various  members  to  the  rotating  cylinders.  The 
exhaust  manifold  is  always  eliminated  and  the  exhaust  gases 
discharge  directly  past  the  valve  to  the  air.  The  carburetor 
(or  equivalent  device)  must  be  stationary  and  discharges  into  the 
crankcase;  separate  induction  pipes  to  the  individual  cylinders 
may  or  may  not  be  provided.  There  is  no  possibility  of  continu- 
ous circulation  of  lubricating  oil;  the  oil  is  used  up  as  it  is  fed. 
The  cylinders  must  be  machined  all  over  to  exact  dimensions  to 
avoid  unbalanced  centrifugal  forces. 

Gnome  Engine. — A  longitudinal  section  of  the  100-h.p.  Gnome 
Monosoupape  (single-valve)  engine  is  shown  in  Fig.  137.  The 
engine  has  nine  cylinders  (each  machined  out  of  a  6-in.  solid 
nickel-steel  bar)  arranged  at  equal  angular  distances  around  the 
crankcase.  The  bore  is  110  mm.,  the  stroke  150  mm.,  clearance 
volume  365  cu.  cm.,  normal  speed  1,200  r.p.m.,  weight  260  Ib. 

The  cycle  of  operations  is  as  follows :  starting  with  any  cylinder 
on  top  center  and  the  exhaust  valve  open,  the  cylinder  draws  in 
air  through  the  exhaust  valve  until  its  closure  at  45  deg.  before  the 
bottom  center;  that  is,  air  is  drawn  in  for  135  deg.  rotation  of 
the  crank.  During  the  next  25  deg.  a  vacuum  is  created  in  the 
cylinder  until,  at  20  deg.  from  the  bottom  center,  the  admission 
ports  are  uncovered  and  a  rich  mixture  enters  from  the  crankcase, 
mixing  with  the  air  already  in  the  cylinder  and  forming  an  explo- 
sive charge.  The  ports  are  again  covered  at  20  deg.  past  bottom 
center  and,  as  the  cylinder  rotates  to  its  top  center,  compression 
occurs.  Ignition  takes  place  at  20  deg.  before  top  center  and 
the  cylinder  rotates  on  its  power  stroke  until  it  is  90  deg.  past 
top  center,  when  the  exhaust  valve  opens  and  remains  open  for 
the  following  405  deg.  rotation.  It  will  be  seen  that  the  exploded 
charge  is  expanded  for  only  half  the  stroke  and  consequently  there 
is  no  possibility  of  high  capacity  or  efficiency  with  this  engine. 
One  important  reason  for  the  early  exhaust  of  the  exploded 
mixture  is  to  prevent  overheating  of  the  cylinders.  It  is  found 
that  a  late  exhaust  will  cause  overheating. 


RADIAL  AND  ROTARY  ENGINES 


181 


182  THE  AIRPLANE  ENGINE 

The  mixture  in  the  crankcase  is  at  all  times  too  rich  to  be 
explosive.  Air  is  drawn  in  through  the  open  rear  end  of  the 
hollow  stationary  crankshaft.  The  fuel  is  supplied  under  air 
pressure  and  sprays  from  the  fuel  nozzle  into  the  crankcase, 
where  it  is  churned  up  with  the  air.  There  is  no  throttle  valve, 
but  the  fuel  supply  can  be  controlled  by  a  regulating  valve  to 
adjust  the  mixture  strength.  The  power  output  of  the  engine  can 
be  varied  through  a  small  range  only,  by  the  use  of  the  regulating 
valve  or  switching  the  ignition  on  and  off.  Variation  of  the 
power  by  cutting  out  the  ignition  on  one  or  more  cylinders  gives 
unequal  impulses,  and  causes  fouling  of  the  cylinders  which  are 
not  firing. 

The  magneto  is  mounted  on  the  face  of  the  back  plate  remote 
from  the  engine.  It  is  driven  through  a  spur  gear  which  meshes 
with  the  large  gear  keyed  to  the  thrust  box  casing.  The  gear 
ratio  is  4  : 9,  that  is,  the  magneto  armature  makes  nine  revolutions 
to  four  of  the  engine,  and  as  the  magneto  gives  two  sparks  per 
revolution  there  will  be  nine  sparks  in  two  revolutions  of  the 
engine.  The  current  from  the  magneto  goes  to  the  distributor 
brush,  which  makes  contact  in  turn  with  the  nine  metal  segments 
of  the  distributor  ring.  The  distributor  ring  revolves  with  the 
engine  and  consequently  18  contacts  are  made  in  two  revolutions 
of  the  engine,  but  no  spark  passes  on  the  exhaust  stroke  as  none 
is  generated  at  the  magneto. 

The  air  pump  (for  the  fuel  pressure)  and  oil  pump  are  mounted 
on  the  back  plate  and  driven  from  the  same  large  gear  as  the 
magneto.  The  oil  pump  delivers  oil  into  two  pipes  of  equal 
size.  Of  the  oil  going  into  one  pipe  about  one-third  flows  through 
a  branch  into  the  thrust  box,  oiling  the  thrust  ball  race  and  main 
engine  ball  race.  The  surplus  oil  overflows  into  the  crankcase 
through  holes  drilled  for  this  purpose,  and  passes  on  to  the 
cylinder  walls  through  the  ports  in  the  base  of  the  cylinder.  The 
main  supply  of  oil  passes  up  the  big  crank  web  through  a  hollow 
plug  in  the  center  of  the  hollow  crankpin,  down  the  short  end 
crank  web  into  the  hollow  short  end  of  the  crankshaft, 
whence  it  is  conveyed  by  a  series  of  holes  to  the  cam  pack.  The 
oil  then  passes  through  grooves  between  the  cams  and  is  thrown 
centrifugally  over  the  interior  of  the  cam  box,  lubricating  the 
cams,  cam  rollers,  tappets,  planet  gear  wheels,  and  the  H  cam 
box  and  nose-piece  ball  races.  The  oil  then  overflows  back  into 
the  crankcase  and  passes  on  to  the  cylinder  walls  as  in  the  case 


RADIAL  AND  ROTARY  ENGINES  183 

of  the  overflow  from  the  thrust  box.  Some  of  the  oil  also  passes 
along  the  hollow  tappet  rods  to  the  rocker  arm  pins. 

The  oil  from  the  other  pipe  flows  up  inside  the  long  end  crank 
web  into  an  annular  space  around  the  brass  plug  in  the  long  end 
crankpin  and  out  of  holes  in  its  balls  to  two  grooves  or  channels 
cut  in  the  ends  of  the  bore  of  the  master  connecting  rod  big  end. 
Holes  are  drilled  from  these  grooves  to  each  wristpin  and  the 
wristpins  are  drilled  to  correspond  with  these  holes,  so  that  the 
oil  may  pass  through  to  lubricate  the  wristpin  bushings.  From 
these  the  oil  passes  into  steel  tubes  (which  are  fixed  to  the  con- 
necting rods)  and  along  the  tubes,  oiling  the  gudgeon  or  wristpins 
and  bushes.  In  later  type  engines  the  steel  tubes  are  dispensed 
with,  and  the  oil  passes  along  the  face  of  the  connecting  rods  to 
the  gudgeon  pins  and  bushes.  The  overflow  from  the  gudgeon 
pins  passes  through  holes  in  the  side  of  the  pistons,  and  lubricates 
the  rings  and  the  cylinder  walls.  The  surplus  oil  is  blown  out 
through  the  exhaust  valve  and  lubricates  the  exhaust  valve-guide 
and  stem. 

The  crankcase  is  made  of  two  steel  stampings  bolted  together 
by  steel  bolts  and  centered  by  dowel  pins.  The  nine  cylinders 
are  each  gripped  tightly  by  the  two  parts  of  the  crankcase  and 
prevented  from  turning  by  a  small  key.  The  crankcase  is 
not  directly  supported  on  the  crankshaft,  but  carries  on  its  faces 
plates  or  covers,  known  respectively  as  the  cam  box  and  the 
thrust  box.  The  thrust  box  contains  the  main  ball  race  and  a 
self-aligning  double-thrust  race.  The  cam  box  contains  the 
planet  gears  and  the  cam  pack  which  actuates  the  exhaust 
valves,  and  one  radial  ball  race.  The  nose  piece  which  carries 
the  propeller  is  mounted  on  the  cam  box. 

The  pistons  are  of  cast  iron  with  concave  heads.  A  portion 
of  the  trailing  edge  is  cut  away  to  allow  the  piston  in  the  adjoin- 
ing cylinder  to  clear.  Each  piston  is  fitted  with  an  obturator 
ring  about  0.6  mm.  thick  in  a  groove  around  its  top.  This 
obturator  ring  is  of  cupped  form  and  is  pressed  out  against  the 
cylinder  wall  by  the  gas  pressure,  thus  preventing  leakage  past 
the  piston.  A  packing  ring  is  fitted  behind  the  obturator  ring 
and  in  the  same  groove.  A  wipe  ring  which  is  made  of  cast 
iron  is  also  fitted  in  a  groove  situated  just  below  the  obturator 
ring.  The  piston  is  fastened  to  its  connecting  rod  by  means  of  a 
hollow  steel  gudgeon  pin  fixed  in  lugs  on  the  underside  of  the 
piston  head  by  means  of  a  tapered  set  screw. 


184  THE  AIRPLANE  ENGINE 

Connecting  Rods. — The  connecting  rod  assembly  consists 
of  a  master  connecting  rod,  to  which  eight  auxiliary  connecting 
rods  are  attached  by  means  of  wristpins.  All  the  rods  are  of 
H-section  and  the  auxiliary  rods  are  bushed  at  both  ends  with 
phosphor-bronze  bushes.  The  master  connecting  rod  big  end  runs 
on  two  ball  bearings;  its  small  end  is  bushed  with  phosphor  bronze. 

The  single  valve  in  the  cylinder  head  performs  the  following 
functions:  (a)  It  acts  as  an  exhaust  valve;  while  so  doing  its 
temperature  is  raised;  (6)  it  admits  to  the  cylinder  a  quantity 
of  air  sufficient  for  combustion  of  the  charge  entering  later 
through  the  ports  at  the  base  of  the  cylinder.  During  this 
portion  of  the  cycle  it  is  cooled. 

The  valve  is  60  mm.  diameter  and  has  a  lift  of  10.5  mm.  It 
is  mounted  in  a  steel  cage  which  also  carries  the  rocker  arm 
fulcrum  pin,  and  is  mechanically  operated  by  means  of  a  hollow 
steel  tappet  rod  and  steel  rocker  arm.  The  valve  stem  slides 
in  a  cast-iron  bush  at  the  center  of  the  cage  which  is  held  in 
position  by  means  of  a  locking  ring  screwed  into  the  cylinder 
head.  The  valve  is  made  heavier  than  is  necessary  for  mechani- 
cal strength  and  is  of  such  weight  as  to  balance  the  centrifugal 
action  of  the  tappet  rod  which  would  otherwise  tend  to  keep  the 
valve  open.  The  valve  spring  is  spiral  and  encircles  the  valve 
stem,  taking  its  bearing  against  the  valve  cage  and  a  detachable 
collar  on  the  valve  stem.  The  valves  are  operated  by  the  cam 
pack,  which  consists  of  nine  cams  keyed  to  a  bronze-bushed  sleeve 
rotating  on  the  small  end  of  the  crankshaft.  The  cams  operate 
the  tappet  rods  which  .Work  the  overhead  rocker  arms.  Each 
tappet  rod  is  formed  of  a  tappet  and  a  rod  jointed  together. 
The  tappet  works  in  a  guide  in  the  cam  box,  and  at  its  inner  end 
is  a  roller  which  bears  against  the  cam.  The  tappet  rod  extends 
from  the  joint  to  the  rocker  arm  of  the  exhaust  valve,  and  is 
adjustable.  The  cam  pack  is  driven  at  half  the  engine  speed  by 
planet  gears,  which  are  fitted  on  the  inner  face  of  the  cover  of  the 
cam  box.  The  engine  is  running  at  twice  the  speed  of  the  cam 
pack,  so  that  the  rollers  at  the  bases  of  the  tappet  rods  are  over- 
taking the  cam  pack.  The  clearance  between  the  rocker  arm 
and  the  bottom  of  the  slot  in  the  valve  stem,  when  the  tappet 
roller  is  at  the  bottom  of  the  cam,  should  be  0.5  mm.  when  the 
engine  is  cold.  In  later  type  engines  the  rocker  arm  engages 
the  valve  stem  by  means  of  a  roller  which  bears  against  the  end 
of  the  stem. 


RADIAL  AND  ROTARY  ENGINES 


185 


Le  Rhone. — The  nine-cylinder  110-h.p.  Le  Rhone  engine 
is  shown  in  Figs.  138  and  139.  The  bore  is  112  mm.,  stroke  170 
mm.,  the  normal  speed  is  1,200  r.p.m.,  weight  323  Ib.  The  80- 
h.p.  nine-cylinder  engine  is  of  105  mm.  bore,  140  mm.  stroke. 
The  cylinder  has  a  cast-iron  liner.  This  engine  differs  from  the 
Gnome  engine  in  several  important  respects.  It  has  an  inlet 
valve  as  well  as  an  exhaust  valve  and  consequently  has  to  be 


Valve  rocker  gear 


A=  Inlet valve 

cam  follower 
B=Ca/77  follower 
lever 

C  "Exhaust  valve 
cam  follower 


D'/n/efcam 
Exhaust  cam 

F=  Internal gear 

mourrkcreccentrica/ly 


FIG.  138.— Transverse  section  of  Le  Rhone  110. 

provided  with  inlet  pipes  from  the  crankcase  to  each  inlet  valve 
cage.  The  valve  timing  is  as  follows:  the  exhaust  valve  closes 
5  deg.  after  top  center  on  the  suction  stroke;  the  inlet  valve  opens 
13  deg.  later  or  18  deg.  after  top  center;  the  inlet  valve  closes 
35  deg.  after  bottom  center  and  compression  goes  on  till  26  deg. 
before  top  center,  when  ignition  occurs.  The  expansion  occupies 
125  deg.  of  the  power  stroke  when  the  exhaust  valve  opens  at 
55  deg.  before  bottom  center  and  remains  open  for  140_deg. 


186 


THE  AIRPLANE  ENGINE 


This  timing  differs   notably  from  that  of  the  Gnome  engine  and 
gives  much  more  complete  expansion  of  the  gases. 

A  rudimentary  carburetor  (see  p.  288)  is  located  at  the  rear 
end  of  the  hollow  crankshaft,  admitting  an  explosive  mixture 
to  the  crankcase.  Control  of  power  output  is  through  the 
throttle  valve.  Ignition  is  by  a  magneto  and  distributor  similar 
in  location  and  general  construction  to  those  of  the  Gnome 


--Rocking  lever  fulcrum 


FIG.  139. — Longitudinal  section  of  Le  Rhone  110. 

engine.     The  pistons  are  convex  and  of  semi-steel.     The  con- 
necting rods  are  of  the  slipper  type  (see  p.  204). 

The  two  valves  on  each  cylinder  are  actuated  by  the  motion 
of  a  rocking  lever  which  is  fulcrumed  at  its  middle  in  ball  bearings. 
This  lever  is  operated  by  a  valve-actuating  rod  which  receives  its 
motion  from  the  trailing  end  of  a  cam-follower  lever.  The  cam- 
follower  lever  is  fulcrumed  at  its  middle  and  carries  an  inlet 
valve  cam  follower  at  the  forward  end  and  an  exhaust  valve  cam 


RADIAL  AND  ROTARY  ENGINES 


187 


follower  at  the  trailing  end.  The  two  cam-followers  are  in 
different  planes  and  are  acutated  by  the  inlet  and  exhaust  cams 
respectively.  These  cams  are  lobed  plates  mounted  on  a  spider 
running  in  ball  bearings  on  a  shaft  eccentric  to  the  crankshaft. 
The  spider  carries  an  internal  gear  into  which  meshes  an  external 
gear  mounted  in  ball  bearings  on  the  crankshaft  and  rotating 
with  the  crankcase.  These  arrangements  are  shown  in  the  right 
half  of  Fig.  138  and  also  in  Fig.  139.  The  external  gear  has  45 
teeth;  the  internal  gear  50  teeth;  consequently  one  complete 
revolution  of  the  engine  will  produce  nine-tenths  of  a  revolution 
of  the  cams.  The  engine  is  overrunning  the  cams  and  would 
100 

90 
80 
70 

Uo 


Q.50 
«> 

0) 

|  40 
30 
20 
10 


Gross  M.E.P. 


95 


90 


75 


70 


800 


1300 


900  1000  1100  1200 

Revolutions  per  Minute 

FIG.  140. — Performance  curves  of  Le  Rhone  80. 

overrun  them  one  complete  revolution  in  10  revolutions  of  the 
engine.  Each  cam  has  five  lobes.  Each  inlet  and  exhaust  valve 
should  be  opened  once  only  in  two  revolutions  of  the  engine,  and 
this  will  be  accomplished  when  the  engine  overruns  the  cam  one- 
fifth  of  a  revolution.  As  the  engine  overruns  the  cams  one- 
tenth  of  a  revolution  each  engine  revolution  it  is  evident  that 
two  revolutions  of  the  engine  are  required  to  complete  the 
opening  and  closing  of  all  the  valves. 

The  action  of  centrifugal  force  on  the  valve-actuating*  rod 
causes  it  to  press  continuously  against  the  valve  rocker  lever. 
At  low  speeds  of  revolution  this  force  may  not  be  sufficient  to 
open  the  exhaust  valve  at  the  desired  time  and  consequently  an 
exhaust  cam  is  necessary  to  push  the  valve  rod  out.  At  high 
speeds  the  operation  of  both  valves  can  be  taken  care  of  by  the 


188 


THE  AIRPLANE  ENGINE 


RADIAL  AND  ROTARY  ENGINES 


189 


inlet  cam  if  it  is  properly  shaped  for  that  purpose.  The  valves 
are  brought  back  to  their  seats  by  spiral  springs  at  low  speeds;  at 
high  speeds  centrifugal  force  closes  the  valves. 

Performance  curves  for  the  80-h.p.  Le  Rhone  are  given  in 
Fig.  140. 

Clerget. — The  Clerget  rotary  engines  are  built  with  7,  9  or  11 
cylinders.  The  130-h.p.,  nine-cylinder  engine  shown  in  longi- 
tudinal section  in  Fig.  141  is  120  mm.  bore,  160mm.  stroke,  makes 


Cam  Gear  Box 


Exhaust  Tappet 
Guide: 


•Inlet  Tappet. 


Locking  Hut. 
•ExhaustCam. 

LockincjNut. 
-Oil  Hole. 

Inlet  Cam. 


Cam6egr 
Cover..--' 


FIG.  142. — Cam  gears  of  B.R.  2. 

1,250  r.p.m.,  weighs  381  lb.,  develops  135  h.p.  and  has  a  compres- 
sion ratio  of  4.  Its  points  of  difference  from  the  previous  engines 
include  the  use  of  an  aluminum  piston,  tubular  connecting  rods, 
inlet  and  exhaust  valves  operated  by  means  of  separate  cams, 
tappets  and  rocker  arms,  and  a  double-thrust  ball  race  which  is  a 
pure  thrust  bearing  and  distinct  from  the  combined  thrust  and 
radial  bearings  of  the  other  engines. 

The  inlet  and  exhaust  cam  plates  are  driven  at  nine-eighths 
the  engine  speed  by  separate  internally-toothed  gears  mounted 


190 


THE  AIRPLANE  ENGINE 


inside  and  keyed  to  the  cam-gear  case.  These  mesh  with  external 
gears  mounted  eccentrically  on  the  crankshaft;  the  cams  are 
attached  to  these  external  gears.  This  arrangement  is  the 
reverse  of  that  used  on  the  Le  Rhone  engine.  The  cam  plates 
overtake  the  engine  once  in  eight  revolutions.  Each  cam  plate 
has  four  lobes  so  that  in  eight  revolutions  each  tappet  will  be 
lifted  four  times,  or  once  in  two  revolutions.  A  sectioned  per- 
spective view  of  the  similar  cam-gear  box  of  the  B.R.2  rotary 
engine  is  shown  in  Fig.  142.  The  four  cams  on  each  gear  are 
simply  rearward  extensions  of  every  fourth  tooth. 

The  valve  timing  differs  in  some  respects  from  that  of  the 
Gnome  and  Le  Rhone.     At  top  center  of  the  suction  stroke 


no 


1150  1200  1250  1300 

Revolu+ions  per    Minute 

FIG.  143. — Performance  curves  of  Clerget  130. 


1350 


both  exhaust  and  inlet  valves  are  open.  The  inlet  opens  5  deg. 
before  top  center;  the  exhaust  closes  5  deg.  past  top  center.  The 
inlet  remains  open  till  58  deg.  past  bottom  center  (or  a  total  of 
153  deg.)  and  compression  begins.  Ignition  is  at  25  deg.  before 
top  center  and  exhaust  begins  68  deg.  before  bottom  center. 
The  carburetor  is  located  at  the  rear  end  of  the  hollow  crank- 
shaft. Fuel  is  injected  under  air  pressure  through  a  jet  which  is 
controlled  by  a  needle  valve.  The  air  supply  is  controlled  by  a 
cylindrical  throttle  valve.  Equal  movement  of  both  throttle 
lever  and  needle  valve  lever  controls  air  supply  only.  Operation 
of  the  throttle  lever  alone  controls  both  air  and  fuel.  The 
charge  -entering  the  crank  passes  to  the  annular  inlet  chamber 
at  the  rear  of  the  crankcase  and  then  by  the  separate  inlet  pipes 
to  the  cylinders.  The  connecting  rod  assembly  is  similar  to  that 
of  the  Gnome  engine  (see  p.  203). 


RADIAL  AND  ROTARY  ENGINES 


191 


The  performance  curve  for  this  engine  (Fig.  143)  is  typical 
of  rotary  engines.  The  effective  horse  power  goes  through  a 
maximum  at  1,250  r.p.m.  but  is  very  flat  for  a  considerable  range 
of  speed.  The  rapidly  increasing  difference  between  the  effective 
horse  power  and  the  indicated  horse  power  is  due  to  the  rapid 
increase  in  the  air-churning  resistance. 

B.R.2  Engine. — The  British  B.R.2  engine  is  one  of  the  largest 
air-cooled  rotary  engines.  In  general  construction  it  is  similar 
to  the  Clerget.  The  cylinders  are  of  aluminum  with  steel  liners 
and  steel  head.  The  cylinder  diameter  is  140  mm.,  stroke  180 
mm.,  compression  ratio  5.01,  brake  horse  power  230  at  1,300 
r.p.m.,  weight  dry  498  lb.,  weight  per  brake  horse  power  dry 


FIG.  144.— Thrust  box  of  B.R.2. 

2.16  lb.  The  cam  gear  for  this  engine  is  shown  in  Fig.  142  and 
follows  exactly  the  same  principle  as  the  Clerget  cam  gear.  The 
thrust  box  contains  two  ball  bearings  and  a  thrust  bearing  which 
differs  from  the  Clerget  in  having  one  row  of  balls  only.  This 
single-thrust  bearing  is  adapted  both  for  pusher  and  tractor 
use  as  indicated  in  Fig.  144;  a  very  small  clearance  is  left  for  the 
travel  of  the  crankcase  along  the  crankshaft  when  changing  from 
pusher  to  tractor. 

Double  Rotary. — The  double-rotary  engine  has  cylinders 
revolving  in  one  direction  while  the  crankshaft  revolves  in  the 
other  direction.  The  effective  speed  is  the  sum  of  the  two  speeds 
so  that  the  power  of  an  engine  in  which  both  cylinders  and  crank- 
shaft revolve  at  900  r.p.m.  is  the  same  as  that  of  a  radial  or 
rotary  engine  of  the  same  dimensions  operating  at  1,800  r.p.m. 
Such  a  speed  is  permissible  in  radial  engines  but  would  give 


192 


THE  AIRPLANE  ENGINE 


excessive  air-churning  resistance  in  a  rotary.  There  is  no  reason 
why  even  higher  effective  speeds — up  to  2,400  r.p.m. — may  not 
be  practicable  with  this  type,  if  the  volumetric  efficiency  of  the 
engine  can  be  maintained  and  if  the  cylinders  can  be  kept  cool 
enough.  In  any  case  this  arrangement  leads  to  a  combination 
of  high  engine  speed  and  low  propeller  speed  with  consequent 
high  propeller  efficiency.  It  has  important  advantages  over  all 
other  types  in  (a)  the  possibility  of  the  elimination  of  unbalanced 


Carburetor 


FIG.  145. — Longitudinal  section  of  Siemens-Halske  double  rotary. 

gyroscopic  effects,  which  is  an  advantage  for  maneuvering,  and 
(&)  the  elimination  of  the  unbalanced  turning  moment  exerted 
by  the  engine  on  the  plane.  This  unbalanced  turning  moment 
is  a  constant,  though  small,  power  drag  on  the  plane  and  its 
elimination  is  a  distinct  advantage. 

The  only  engine  of  this  type  which  has  been  in  production  is 
the  Siemens-Halske  11-cylinder  engine  (Fig.  145),  which  was 
brought  out  in  1918  and  develops  200  h.p.  at  900  r.p.m.  of  both 
cylinders  and  crankshaft,  or  a  virtual  speed  of  1,800  r.p.m. 


RADIAL  AND  ROTARY  ENGINES  193 

The  single  propeller  is  mounted  on  a  nose  attached  to  the  revolv- 
ing crankcase;  the  torque  of  the  crankshaft  is  transmitted  to  the 
crankcase  by  securing  a  bevel  wheel  to  the  crankshaft  and  a 
similar  gear,  facing  it,  to  the  crankcase  and  mounting  an  inter- 
mediate pinion  between  the  two  on  a  stud  which  is  fastened  to 
the  stationary  cylindrical  housing  at  the  rear  of  the  engine. 

The  carburetor  is  mounted  on  a  stationary  hollow  extension 
of  the  crankshaft.  The  combustible  charge  is  drawn  in  through 
the  hollowcrank  shaft  to  the  crankcase  and  goes  from  the  annular 
inlet  chamber  at  the  rear  of  the  crankcase  to  the  individual  inlet 
pipes.  The  inlet  and  exhaust  valves  are  operated  through  two 
cam  plates  which  are  loose  on  the  crankshaft  and  are  rotated 
through  double  reduction  gears  from  an  internal  gear  attached 
to  the  crankcasing.  The  engine  is  supported  by  steel  rods  both 
before  and  behind  the  cylinders. 

The  weight  of  the  engine  complete  is  427  lb.,  which  at  240 
maximum  h.p.  gives  a  weight  of  1.78  lb.  per  horse  power.  The 
fuel  economy  is  as  good  as  with  stationary  engines  and  is  much 
better  than  with  other  rotaries. 

Other  designs  of  double  rotaries  with  two  propellers  (right- 
and  left-hand  respectively)  forward  of  the  cylinders,  attached 
to  the  crankshaft  and  crankcase  respectively,  have  not  passed  the 
experimental  stage.  The  efficiency  of  a  pair  of  propellers  close 
together  but  operating  in  opposite  directions  has  been  found  to 
be  but  little  inferior  to  that  of  a  single  propeller.  There  is 
consequently  the  possibility  of  the  development  of  a  satisfactory 
double-rotary  engine  on  the  lines  indicated. 

Radial  Engines. — In  a  fixed  radial  engine  the  cylinders  are 
stationary  and  the  crankshaft  revolves.  Three  to  eleven  cylin- 
ders can  be  accommodated  in  a  single  row  or  bank,  but  two  rows 
with  a  two-throw  crank  must  be  adopted  if  a  larger  number  of 
cylinders  is  desired  or  if  it  is  necessary  to  cut  down  the  over-all 
diameter.  The  two-throw  crank  eliminates  the  need  for  counter- 
balance weights  but  increases  the  length  of  the  engine  and  intro- 
duces difficulties  in  the  air  cooling  of  the  rear  row  cylinders. 

Radial  engines  offer  certain  special  construction  problems. 
The  most  important  are  the  balancing  of  the  masses  at  the  crank- 
pin  and  the  avoidance  of  excessive  pressures  on  the  crankpin. 
It  is  possible  to  operate  radial  engines  at  speeds  as  high  as  those 
used  in  vertical  and  Vee  engines,  but  special  care  must  be  taken 

to  prevent  the  overheating  of  air-cooled  cylinders  and  the  over- 
is 


194 


THE  AIRPLANE  ENGINE 


loading  of  the  crankpin.  There  is  no  fundamental  reason  why 
the  mean  effective  pressure  and  economy  of  radial  engines  should 
not  be  as  good  as  those  of  any  other  type. 

A  B  C. — The  development  of  air-cooled  fixed  radial  engines 
has  been   carried  on  in   England  more  -than  elsewhere.     The 


FIG.  146. — Sectional  outlines  of  A  B  C  "  Dragonfly."     Radial  engine. 

ABC  engines,  built  by  the  Walton  Motors  Co.,  have  the  following 
general  characteristics : 


Type  name 

Gnat 

Wasp 

Dragonfly 

Number  of  cylinders  (copper-coated 
steel  fins)   . 

2 

7 

9 

Bore,  inches  

4  75 

4  75 

5  5 

Stroke,  inches 

5  5 

6  25 

6  5 

Normal  brake  horse  power  
Revolutions  per  minute  
Oil  consumption,  pints  per  hour  .  . 
Gasoline   per   brake    horse    power 
hour,  pints 

45 
1,800 
1.7 

0  56 

.       200 
1,800 
4 

0  56 

340 
1,650 

7 

0  56 

Weight  of  engine,  dry,  pounds  

115 

320 

600 

Weight    per    brake    horse    power, 
pounds  

2.3 

1.6 

1  765 

Over-all  diameter,  inches 

42  7 

50  5 

RADIAL  AND  ROTARY  ENGINES 


195 


The  Wasp  and  Dragonfly  engines  have  each  two  exhaust  valves 
and  one  inlet  valve  per  cylinder.  Their  engines  use  the  master- 
rod  connecting-rod  assembly  with  roller  bearings  (see  p.  207)  and 
have  counterbalance  weights.  Sectional  views  of  the  Dragonfly 
engine  are  shown  in  Fig.  146. 

Cosmos. — The  Cosmos  Engineering  Co.  has  fixed  radial  engines 
with  the  following  characteristics: 


Type  name 

Lucifer 

Jupiter, 
direct 
drive 

Jupiter, 
geared 

Mercury 

Hercules, 
geared 

Number  of  cylinders  .  .  . 
Number  of  rows  

3 
1 

9 
1 

9 
1 

14 
2 

18 
2 

Bore,  inches  

5.75 

5.75 

5.75 

4  375 

6  25 

Stroke,  inches  
Normal      brake      horse 
power  

6.25 
100 

7.5 
400 

7.5 
450 

5% 
315 

7.5 
1,000 

Brake     mean    effective 
pressure,    pounds    per 
square  inch 

113 

113 

Revolutions  per  minute. 
Propeller  speed,  revolu- 
tins  per  minute 

1,600 

1,650 

1,850 
1  200 

1,800 

1,750 
1  150 

Weight  of  engine,  dry, 
pounds 

220 

636 

757 

587 

1  400 

Weight  per  brake  horse 
power,  pounds   

2.2 

1.59 

1  863 

1  4 

Weight  per  brake  horse 
power    at    maximum 
power,  pounds 

1  413 

Over-all  diameter,  inch. 

52.5 

52.5 

41.625 

The  Jupiter  engine  has  two  exhaust  and  two  inlet  valves;  the 
Mercury  engine  has  two  exhaust  and  one  inlet  valve. 

Performance  curves  for  the  Jupiter  engine  are  shown  in  Fig. 
147.  It  will  be  seen  that  the  brake  mean  effective  pressure 
reaches  a  maximum  of  117  Ib.  per  square  inch  at  1,700  r.p.m. 

A  special  feature  of  the  Jupiter  engine  is  the  method  of  con- 
veying the  explosive  charge  to  the  cylinders.  There  are  three 
independent  carburetors  at  the  rear  of  the  engine  discharging 
into  the  cover  of  the  annular  inlet  chamber  which  forms  the  rear 
of  the  crankcase.  This  chamber  (Fig.  148)  contains  an  alu- 
minum spiral  casting  which  fits  closely  into  the  chamber.  The 


196 


THE  AIRPLANE  ENGINE 


560 
540 
520 
500 
480 
460 
440 
420 
^400 

1380 

cti 
360 

340 
320 
300 
280 
260 
240 
220 
200 

147  — 

s£ 

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DOO       1200        1400         1600          1800        2000      2200 
R.P.  M. 

Performance  curves  of  Cosmos  "Jupiter"  radial  engine. 

2     ^(^S—       3 
FIG.  148. — Induction  chamber  of  Cosmos  "Jupiter.' 


RADIAL  AND  ROTARY  ENGINES 


197 


casting  constitutes  a  three-part  spiral.     The  carburetors  dis- 
charge into  the  spaces  marked  X,  Y  and  Z  respectively.     The 


FIG.  149. — Longitudinal  section  of  Salmson  radial  engine. 

space  X  is  part  of  the  spiral  marked  AAA,  so  that  the  mixture 
drawn  into  X  will  flow  along  the  spiral  groove  AAA.     This 


198 


THE  AIRPLANE  ENGINE 


groove  is  opposite  the  inlet  pipes  for  cylinders  2,  8  and  5;  similarly 
the  middle  carburetor  will  supply  cylinders  3,  9  and  6.  This 
arrangement  gives  the  mixture  a  clean  sweep  from  the  carburetor 
to  the  cylinder  and  isolates  the  cylinders  in  three  groups  so  that 
should  one  carburetor  fail  to  act  properly  there  would  still  be  six 
cylinders  in  normal  action. 


Exhaust  Valve 


---fn/ef  Valve 


FIG.  150. — Transverse  view  of  Salmson  radial  engine. 

Salmson. — The  Salmson  (Canton-IInne*)  engine  is  a  good 
example  of  the  water-cooled  fixed-radial  engine.  Figure  149  shows 
a  longitudinal  section  of  a  nine-cylinder  engine;  Fig.  150  is  a 
transverse  view  of  the  same  engine.  The  general  dimensions 
of  the  engine  are:  bore,  125  mm.;  stroke,  170  mm.;  ratio  of 
compression,  5.3;  weight  of  engine  without  water  or  radiator, 


RADIAL  AND  ROTARY  ENGINES 


199 


474  lb.;  weight  of  water  in  jackets  20  lb.;  power  at  1,500  r.p.m., 
250  h.p.;  weight  dry  per  horse  power,  1.89  lb.;  gasoline  con- 
sumption per  horse-power  hour,  0.507  lb.;  oil  consumption  per 
horse-power  hour,  0.077  lb.  The  variation  of  brake  horse  power 
with  engine  speed  is  shown  in  Fig.  151. 

The  cylinders  are  steel  forgings  3  mm.  thick;  the  jackets 
are  of  sheet  steel  welded  to  the  cylinders.  The  inlet  and  exhaust 
valves  are  symmetrically  located  and  are  both  62.5  mm.  dia- 
meter; they  are  held  to  their  seats  by  rat-trap  springs.  The 
connecting-rod  assembly  is  of  the  master-rod  type  (see  p.  203) 
with  ball  bearings  on  the  crankpin.  The  crankshaft  is  of  the 


300 


Brake  Horsepower 

ro  ro  r  ro  I 

8  S  si  J 

s* 

" 

s^ 

^ 

s 

s* 

s 

/ 

^ 

/ 

" 

S 

/ 

s 

/ 

/ 

/ 

/ 

S 

^/ 

/ 

1300                      1400                       1500                      1600                     IKX 
Revolutions  per  Minute 

FIG.  151. — Performance  curve  of  Salmson  radial  engine. 

built-up  type  with  counterweights.  The  valves  are  operated 
through  push  rods  and  rocker  arms  from  a  cam  sleeve  which  is 
revolved  on  the  crankshaft  at  one-fourth  the  engine  speed  by 
means  of  an  epicyclic  gear  set.  There  are  three  pairs  of  cams  on 
the  cam  sleeve,  each  pair  at  opposite  diameters  in  its  own  plane. 
In  each  of  the  three  planes  are  the  cam  followers  of  both  valves 
for  three  cylinders.  All  the  valves  will  be  opened  twice  in  one 
revolution  of  the  cam  sleeve.  Consequently  in  two  revolutions 
of  the  engine,  or  one-half  revolution  of  the  cam  sleeve,  each  of  the 
valves  will  have  been  operated  once. 

The  inlet  valve  opens  at  top  center  and  closes  55  deg.  after  the 
bottom  center,  or  at  about  16  per  cent  of  the^re turn  stroke. 
Ignition  occurs  about  30  deg.  before  top  center.  The  exhaust 
opens  65  deg.  before  bottom  center  and  closes  at  top  center. 


200 


THE  AIRPLANE  ENGINE 


The  water  circulation  is  shown  in  Fig.  152.  A  centrifugal 
pump  taking  water  from  the  bottom  of  the  radiator  discharges 
it  through  two  pipes  into  the  heads  of  the  two  lowest  cylinders. 
The  top  and  bottom  of  each  jacket  is  connected  by  pipes  to  the 
tops  and  bottoms  respectively  of  the  adjacent  cylinders.  The 
water  is  finally  delivered  from  the  top  of  the  highest  cylinder  to 
the  radiator. 

The  carburetor  (Zenith)  discharges  through  long  vertical  pipes 
into  the  annular  inlet  chamber  at  the  rear  of  the  crankcase  and 
thence  through  separate  inlet  pipes  to  the  individual  cylinders. 


Wafer  Screen — . 


Thermometer, 


-Radiator 


FIG.  152. — Water  circulation  in  Salmson  radial  engine. 

The  exhaust  passes  from  each  cylinder  into  a  sheet  metal  exhaust 
duct  which  encircles  the  engine,  discharges  at  the  sides  of  the 
fuselage,  and  is  stream-lined  to  serve  as  a  cowling  for  the  engine. 
Details  of  Radial  and  Rotary  Engines. — Air-cooled  cylinders 
are  either  made  from  solid  steel,  as  in  the  Gnome,  Le  Rhone  and 
Clerget  rotary  engines,  or  they  are  composite  with  steel  barrel 
or  liner  and  aluminum  alloy  head.  All-aluminum  cylinders  have 
been  tried  with  fair  success  but  there  is  doubt  of  their  durability; 
they  are  no  lighter  than  the  other  types  and  their  considerable 
longitudinal  expansion  increases  tappet  clearances  and  alters 
valve  timing  to  a  greater  extent  than  with  other  constructions. 


RADIAL  AND  ROTARY  ENGINES 


201 


EXHAUST 


The  satisfactory  operation  of  an  air-cooled  cylinder  depends 
on  keeping  down  its  temperature.  When  overhead  valves  are 
used,  this  temperature  is  highest  in  the  middle  of  the  head. 
With  open  exhaust  as  in  rotary  engines  and  in  some  radial  engines 
there  is  not  much  difficulty  in  arranging  for  adequate  cooling  of 
all  parts  of  the  cylinder. 

With  overhead  valves  it  is  essential 
to  make  the  cylinder  head  of  the 
best  available  conductor  (see  p.  346), 
which  in  practice  turns  out  to  be  an 
aluminum-copper  alloy.  The  valve 
seats  and  the  working  surface  of  the 
cylinder  barrel  must  be  of  some  harder 
material.  When  an  aluminum  head 
is  used  the  valve  seats  should  consist 
of  rings  of  steel  or  bronze,  cast  or 
expanded  into  position.  Bronze  seats, 
in  consequence  of  their  high  coefficient 
of  expansion,  are  less  likely  to  come 
loose  than  steel  seats. 

One  type  of  construction  is  shown 
in  Fig.  153.  An  aluminum  casting 
forms  the  head  and  surrounds  the 
greater  part  of  the  steel  liner,  which 
is  shrunk  into  the  casting  at  about 
300°C.  Cylinders  of  this  type  have 
given  excellent  results,  but  the  differ- 
ence between  the  coefficients  of  ex- 
pansion Of  the  Steel  and  aluminum  FIG- 153.— Aluminum  air-cooled 

cylinder  with  steel  liner. 

tends  to  cause  separation  of  the  liner 

and  casing  at  working  temperatures  and  a  film  of  oil  may  work  in 
between  them.  With  cylinders  below  4  in.  in  diameter  there 
is  little  trouble.  A  shrinkage  allowance  of  about  1  in  600 
should  be  made.  The  expansion  trouble  can  be  overcome  by  the 
use  of  bronze  liners  if  a  bronze  sufficiently  hard  to  resist  wear 
is  developed.  The  holding-down  bolts  go  through  lugs  in  the 
aluminum  casting. 

Screwed-in  liners  have  not  given  good  results  owing  to  the 
impossibility  of  maintaining  adequate  contact  between  liner 
and  casing.  If  the  contact  is  good  when  cold,  the  difference  of 
expansion  when  hot  causes  contact  at  points  only. 


202 


THE  AIRPLANE  ENGINE 


The  best  method  of  composite  construction  is  one  with  an 
aluminum  cylinder  head  into  which  is  cast  or  screwed  a  steel 
barrel  with  its  own  cooling  fins  (Fig.  154).  This  construction  is 
mechanically  sound  and  has  been  used  successfully  with  cast-in 
barrels  for  sizes  up  to  6  in.  in  diameter  and  with  screwed-in  barrels 
up  to  5^  in.  in  diameter.  The  length  of  the  screwed  portion 
should  be  about  one-fourth  of  the  cylinder  diameter.  The 

holding-down  bolts  grip  a  ring  integral 
with  the  steel  barrel  and  thereby  avoid 
the  breakages  of  holding-down  lugs 
which  have  been  rather  frequent  with 
the  construction  of  Fig.  153.  In  an- 
other type  of  construction  the  barrel 
and  head  are  formed  ^of  steel  in  one 
piece  and  an  aluminim  cap  embody- 
ing the  inlet  and  exhaust  ports  is 
bolted  to  the  cylinder  head. 

Tests  of  all-steel  cylinders  such  as 
are  used  in  Le  Rhone  and  Clerget 
engines,  with  cylinder  diameters  rang- 
ing from  4  to  6  in.,  show  that  the  all- 
steel  cylinder  gives  very  appreciably 
higher  fuel  consumption  and  lower 
brake  mean  effective  power  than 
does  the  aluminum-headed  cylinder.1 
A  5M  by  6H-m-  steel  cylinder  with 
one  aluminum  inlet  and  two  cast-iron 
exhaust  ports  bolted  to  it  was 
changed  (1)  by  having  an  aluminum  cap  bolted  to  its  head  and 
(2)  by  having  the  original  head  cut  off  and  an  aluminum 
head  cast  on  to  the  same  barrel.  Tests  showed  that  under 
maximum  load  conditions  at  1,450  r.p.m.  and  in  a  wind  of  82 
miles  per  hour  the  aluminum  headed  cylinder  gave  15  per  cent 
more  power  than  either  of  the  others.  The  fuel  consumption 
was  26  per  cent  less  than  that  of  the  steel  cylinder  and  20  per  cent 
less  than  that  of  the  capped  cylinder. 

A  capped  steel  cylinder  is  usually  not  much  better  than  the 
normal  steel  cylinder;  however  well  fitted  initially,  " growth" 
and  distortion  of  the  aluminum  impair  the  contact  after  a  few 
hours'  running. 

1  A.  H.  GIBSON,  Inst.  Aut.  Eng.,  Feb.,  1920. 


FIG.  154. — Steel  air-cooled  cyl- 
inder with  aluminum  head. 


RADIAL  AND  ROTARY  ENGINES  203 

The  largest  all-steel  air-cooled  cylinder  tested  by  Gibson  was 
6  by  8  in.  With  a  compression  ratio  of  4.48  and  in  a  wind  of 
75  miles  per  hour  this  cylinder  developed  115  Ib.  brake  mean 
effective  pressure  on  a  fuel  consumption  of  0.68  Ib.  per  brake- 
horse-power  hour  at  1,250  r.p.m.,  and  105  Ib.  brake  mean  effective 
pressure  at  1,600  r.p.m.  An  aluminum-headed  cylinder  of  the 
same  dimensions  developed  under  the  same  conditions  121  Ib. 
brake  mean  effective  pressure  on  a  consumption  of  0.56  Ib.  per 
brake  horse  power  per  hour. 

Cylinder  distortion  may  arise  from  the  fact  that  the  cooling 
air  blast  is  directed  against  one  side  of  the  cylinder.  Such 
distortion  is  negligible  when  the  blast  is  directed  on  the  exhaust 
side.  This  side  is  normally  the  hottest  and  needs  most  cooling. 
Tests  on  a  5^-in.  aluminum  cylinder  with  the  blast  on  the 
exhaust  side  showed  a  maximum  temperature  difference  between 
the  front  and  back  of  the  barrel  of  58°C.,  and  a  mean  difference 
of  19°C.  With  the  blast  on  the  inlet  side  the  maximum  tem- 
perature difference  was  180°C.  and  the  mean  120°C.  In  spite  of 
this  the  cylinder,  which  was  fitted  with  an  aluminum  piston  of 
only  0.025-in.  clearance,  gave  no  sign  of  binding,  showing  that 
even  in  this  extreme  case  the  distortion  was  not  serious. 

With  longitudinal  fins  and  a  comparatively  uniform  distribu- 
tion of  air  flow,  the  distortion  is  not  noticeably  less.  The 
exhaust  side  will  be  the  hottest  and  the  temperature  will  be  less 
uniform  than  with  circumferential  fins  and  a  free  blast  on  the 
exhaust  side.  Furthermore,  longitudinal  fins  do  not  stiffen  the 
cylinder  as  strongly  against  distortion  as  do  circumferential  fins. 

Connecting-rod  Assembly. — The  problem  of  connecting  seven 
or  nine  big-ends  to  a  single  crankpin  is  usually  solved  either  by 
the  "articulated  or  master  rod"  assembly  or  by  the  "slipper" 
assembly. 

The  master  rod  assembly  is  used  on  the  Gnome  and  Clerget 
rotaries  and  on  most  of  the  radials.  Details  of  the  assembly,  as 
installed  in  the  Salmson  engine,  are  given  in  Figs.  155  and  156. 
The  big  end  of  the  master  rod  encircles  the  crankpin,  holds  the 
wristpins  for  all  the  short  rods,  and  carries  the  outer  races  of  the 
ball  bearings.  It  will  be  seen  that  this  construction  shortens  the 
effective  length  of  all  rods  except  the  master  rod;  that  the  axes 
of  the  short  rods  pass  through  the  crankpin  only  twice  in  the 
revolution;  and  that  the  obliquity  of  the  short  rods  is  considerably 
greater  than  that  of  the  master  rod. 


204 


THE  AIRPLANE  ENGINE 


The  slipper  type  of  assembly  is  used  in  the  Le  Rhone  and 
Anzani  engines.     The  crankpin  carries  on  ball  bearings  (Figs. 


FIG.  155. — Articulated  connecting-rod  assembly. 

157  and  158)  two  thrust  blocks  each  of  which  has  three  annular 
grooves  lined  with  bearing  metal.     The  two  discs  are  fastened 


^ 


FIG.  156. — Section  through  articulated  connecting-rod  assembly. 

together  with  the  annular  grooves  opposite  one  another.     The 
big  ends  of  the  nine  connecting  rods  are  provided  with  slippers 


RADIAL  AND  ROTARY  ENGINES 


205 


\ 


'IG.    157. — Section  through  slipper       FIG.  158. — Assembly  of  slipper  type  connect- 
type  connecting-rod  assembly.  ing  rods. 


FIG.  159. — Diagram  of  rotary  engine  with  slipper'type  connecting-rod  assembly. 


206 


THE  AIRPLANE  ENGINE 


each  of  which  is  turned  with  the  same  radius  of  curvature  as  one 
of  the  annular  grooves.  Three  connecting  rods  act  on  each 
groove  and  consequently  there  are  three  designs  of  slipper.  The 
slippers  for  the  middle  and  outermost  grooves  are  slotted  to 
avoid  contact  with  the  connecting  rods  for  the  innermost  and 
middle  grooves.  The  arrangement  is  shown  in  outline  in  Fig. 
159.  The  plan  of  the  slippers  in  Fig.  160  shows  the  slotting  to 
prevent  interference  with  adjacent  connecting  rods. 

The  slipper  assembly  is  considerably  heavier  than  the  master- 
rod  type  and  consequently  is  better  adapted  to  rotaries  than  to 
radials.  It  has  the  advantage  that  the  connecting  rod  is  of 

maximum  length  and  conse- 
quently of  minimum  angularity 
and  also  that  the  thrust  (or  ten- 
sion) of  the  rod  always  passes 
through  the  center  of  the  crank- 
pin.  Furthermore  a  large  bear- 
ing surface  is  provided  at  the 
thrust  block  which  is  easily 

FIG.   160.-Pro.jected   views  of  slip-     lubricated   by  the    oil  thrown  off 

from  the  ball  bearings. 

Dynamical  Comparison  of  Radial  and  Rotary  Engines.— 
The  fixed-radial  engine  presents  the  special  problem  of  a  large 
mass  rotating  with  the  crankpin  and  consequently  large  centrif- 
ugal force.  The  inertia  forces  of  the  reciprocating  parts  are 
additive  to  this.  The  result  is  a  considerable  total  pressure 
on  the  crankpin,  which  is  relieved  somewhat  by  the  gas  pres- 
sures during  the  explosion  strokes.  Roller  or  ball  bearings  are 
necessary  at  the  crankpin  if  high  speeds  of  rotation  are  to  be 
maintained. 

Balancing  of  the  primary  inertia  forces  of  a  single-crank  fixed- 
radial  engine  is  readily  effected  by  a  mass,  approximately  equal  to 
half  the  mass  of  all  the  reciprocating  parts,  used  as  a  counter- 
balance opposite  the  crankpin  at  crankpin  radius.  The  counter- 
balance weight  will  add  7  to  10  per  cent  to  the  weight  of  the 
engine  and  can  be  avoided  only  by  using  two  rows  of  cylinders 
and  a  double-throw  crank.  In  the  last  case  there  is  an  unbal- 
anced primary  couple.  Balancing  the  centrifugal  and  inertia 
pressures  on  the  crankpin  has  been  accomplished  in  an  ingenious 
manner  in  the  latest  design  of  Cosmos  " Jupiter"  engine.  Two 
bob- weights  are  suspended  on  the  outer  sides  of  the  master  rod; 


RADIAL  AND  ROTARY  ENGINES 


207 


their  other  ends  are  connected  to  the  main  crankshaft  balance 
weights  through  hardened  blocks  working  in  slots  machined  in 
the  bob- weights.  The  bob- weights  serve  not  only  to  relieve  the 
pressure  on  the  crankpin  but  also  as  part  of  the  weight  necessary 
to  balance  the  engine  as  a  whole.  The  general  arrangement  of 
these  bob- weights  is  shown  in  Fig.  161. 

In  rotary  engines  the  pistons  and  connecting  rods  rotate  about 
a  stationary  crankpin  with  an  angular  velocity  which  is  variable. 
With  the  master-rod  type  of  connecting-rod  assembly  there  is 


FIG.   161. — Balanced  connecting  rod  of  Cosmos  "Jupiter"  engine. 

some  lack  of  centrifugal  balance  at  the  crankpin,  but  it  is  usually 
negligible.  The  connecting  rods  are  subjected  to  centrifugal 
tensional  loading;  the  pressures  on  the  pins  at  the  ends  of  the 
rods  increase  as  the  square  of  the  revolutions  per  minute.  With 
the  same  connecting  rod  loading  a  fixed-radial  engine  may  run  at 
approximately  twice  the  speed  of  a  rotary  engine  with  the  same 
moving  parts;  before  that  speed  is  reached,  however,  the  crankpin 
loading  of  the  radial  becomes  excessive. 

Ball  and  roller  bearings  for  crankpins  of  radial  engines  offer 
special  problems.     The  bearing  rotates  as  a  whole  and  presents 


208  THE  AIRPLANE  ENGINE 

conditions  of  loading  quite  unlike  those  of  stationary  bearings. 
Considerable  investigation  of  this  matter  was  made  by  the 
British  Department  of  Aircraft  Production1  as  a  result  of  the 
failure  of  both  caged  and  uncaged  bearings. 

An  analysis  of  the  situation  showed  that  with  cageless  bearings 
the  balls  are  crowded  away  from  the  center  of  rotation,  by  centrif- 
ugal force,  and  rub  against  one  another.  The  balls  rotate 
usually  at  about  2,500  r.p.m.  about  their  own  centers;  the  points 
that  touch  are  always  moving  in  opposite  directions  and  the 
abrasion  is  considerable.  When  a  cage  is  used  the  centrifugal 
force  on  the  cage  and  the  balls  causes  a  displacement  of  the  cage 
until  the  bearing  load  on  the  balls  nearest  the  crank  center  due  to 
the  cage  wedging  between  them  is  equal  to  the  total  centrifugal 
load  on  the  cage.  This  causes  a  heavy  abrasive  action  between 
the  balls  and  the  cage. 

For  successful  operation  it  is  necessary  to  have  a  cage  which 
will  carry  independently  the  rubbing  loads  on  each  ball  due  to 
centrifugal  force.  To  accomplish  this  (1)  the  cage  must  be 
strong  enough  to  take  the  independent  loads  from  the  balls 
without  distortion,  (2)  sufficient  bearing  surface  must  be  provided 
at  the  surface  of  location  of  the  cage  to  carry  safely  the  total 
centrifugal  load,  (3)  sufficient  bearing  surface  must  be  provided 
between  the  balls  and  the  cage  to  prevent  wear  on  the  cage,  (4) 
the  cage  must  be  made  of  a  metal  of  minimum  abrasion,  and  (5) 
all  surfaces  must  run  with  a  continuous  flow  of  oil. 

To  meet  these  requirements  a  cage  as  in  Fig.  162  may  be 
used.  This  type  of  cage  must  be  definitely  located  and  not 
displaceable  by  centrifugal  force  for  more  than  a  few  thousandths 
of  an  inch.  The  design  shown  is  made  in  two  halves  with  eight 
hemispherical  holes  with  0.10  in.  clearance  for  the  balls.  The 
outside  circumference  is  turned  in  a  flat  V  to  avoid  the  actual  ball 
path,  and  on  either  side  of  the  V  is  a  true  cylindrical  surface 
about  %Q  in.  wide.  The  cage  is  of  phosphor  bronze  and  fits 
the  outer  ball  race  with  a  clearance  of  0.005  to  0.007  in.  This 
bearing  proved  entirely  satisfactory  on  the  crankpin  of  a  10- 
cylinder,  115  by  150-mm.  radial  Anzani  engine  developing  150 
h.p.  at  1,300  r.p.m. 

The  details  of  a  satisfactorily  located  cage  for  rollers  for  the 
crankpin  for  a  320-h.p.  nine-cylinder  radial  engine  making  1,700 

1  J.  B.  SWAN,  The  Automobile  Engineer,  July,  1919. 


RADIAL  AND  ROTARY  ENGINES 


209 


r.p.m.  are  given  in  Fig.  163.  The  cage  is  located  on  the  roller 
track,  which  in  practice  works  out  advantageously  in  polishing 
the  track  and  keeping  it  free  from  foreign  matter. 


fT 


&*&#***?*' 


view  of  R.H.HQIF 


Sec-Kon  A-A-A 

FIG.  162.— Located  cage  for  ball  bearings. 

Details  of  successful  and  unsuccessful  ball  and  roller  bearings 
are  given  in  the  following  table.  At  speeds  above  1,600  r.p.m. 
and  with  a  radius  of  rotation  above  2*^  in.  cageless  bear- 


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FIG.  163. — Located  cage  for  roller  bearings. 

ings  will  not  run  satisfactorily  if  the  balls  or  rollers  are  larger 
than  %  in.  in  diameter  and  the  inner  race  larger  than  1.5  in.  in 
diameter. 

14 


210 


AIRPLANE  ENGINE 


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RADIAL  AND  ROTARY  ENGINES  211 

The  crankshaft  in  a  radial  engine  will  be  solid  or  built-up 
according  as  the  big  end  of  the  connecting  rod  has  a  plain  bearing 
or  a  ball  or  roller  bearing.  The  plain  bearing  is  likely  to  give 
trouble  in  view  of  the  heavy  loading  of  the  bearing  except  for 
low-speed  engines,  and  can  be  used  only  with  high-pressure 
forced  lubrication;  ball  or  roller  bearings  are  very  generally  used. 

The  built-up  crank  necessary  with  ball  or  roller  bearings  may 
be  either  a  two-web  shaft  with  equal  loading  on  front  and  rear 
bearings  or  an  overhung  crank  with  a  drag  crank  for  driving 
auxiliaries.  The  former  type  is  the  more  desirable.  With  the 
overhung  crank  the  main  balance  weight  is  on  a  single-crank  web, 
which  leads  to  an  unbalanced  couple,  and  also,  the  diameter  of 
shaft  and  bearings  has  to  be  made  greater. 

Valve  Operation. — Apart  from  the  design  of  multilobed  cams 
and  their  driving  gears,  the  valve  operation  of  radial  engines  does 
not  present  any  special  problems.  In  rotary  engines  the  effects 
of  centrifugal  force  on  the  push  rods  and  tappets  have  to  be  met. 
Counterweights  have  sometimes  been  used  on  the  valve  side  of 
the  rocker  arm,  but  their  use  increases  the  load  and  wear  on  the 
cam  profile.  The  necessity  for  keeping  down  the  over-all  engine 
diameter  is  likely  to  result  in  the  selection  of  an  unfavorable  type 
of  valve  spring  and  an  undesirable  reduction  in  the  length  of  the 
valve  stem  guide;  failures  have  been  frequent  when  volute  valve 
springs  have  been  used. 

Lubrication. — Rotary  engines  are  always  wasteful  of  oil, 
using  almost  Ho  lb.  per  brake-horse-power  hour.  There  is  no 
return  of  surplus  oil  to  the  pump,  which  consequently  has  to 
determine  the  amount  of  oil  used.  Plunger  pumps  are  always 
used  discharging  directly  to  the  main  bearings,  big  end  and  cam 
gear,  and  relying  largely  on  centrifugal  force  for  the  lubrication 
of  wristpins  and  cylinders. 

In  radial  engines  either  a  plunger  pump 'or  a  gear  pump  may 
be  used  with  a  dry  sump.  There  is  danger  of  over-oiling  the 
lower  cylinders.  Most  of  the  oil  goes  direct  to  the  crankpin  and 
is  distributed  thence  by  centrifugal  force  to  the  bearings,  connect- 
ing-rod assembly,  and  cylinders.  The  oil  consumption  of  radial 
engines  runs  from  about  0.02  to  0.04  Ib.  per  brake-horse-power 
hour. 


CHAPTER  IX 
FUELS  AND  EXPLOSIVE  MIXTURES 

The  properties  desired  in  an  airplane  engine  fuel  are  as  follows : 

1.  It  must  have  a  high  heat  of  combustion  per  pound.     This 
determines  the  cruising  radius  for  a  given  weight  of  fuel,  since 
efficiency  does  not  vary  appreciably  with  the  fuel. 

2.  It  must  have  a  high  heat  of  combustion  per  cubic  foot  of 
explosive  mixture  if  it  is  to  develop  high  horse  power  per  cubic 
foot  of  piston  displacement.     Alcohol  and  gasoline  have  about 
the  same  heats  of  combustion  per  cubic  foot  of  explosive  mixture 
but  very  different  heats  of  combustion  per  pound  of  fuel.     The 
heat  of  combustion  is  nearly  constant  for  all  the  available  fuels. 

3.  It  must  be  able  to  withstand  high  compression  without 
preignition  or  detonation. 

4.  It  must  vaporize  readily  (preferably  with  little  or  no  pre- 
heating of  the  air)  upon  admixture  with  air  and  should  be  com- 
pletely   vaporized    at   the   beginning   of  explosion.     For  good 
distribution  it  should  be  completely  vaporized  upon  reaching 
the  admission  manifold,  but  this  is  not  usually  attained. 

5.  Combustion  should  be  complete,  leaving  no  solid  residue  in 
the  cylinder. 

6.  The  fuel  and  products  of  combustion  must  not  be  corrosive. 

7.  The  explosion  rate  must  be  neither  too  rapid   (as  with 
hydrogen,  acetylene  and  ether)  nor  too  slow. 

8.  The  bulk  of  fuel  and  the  weight  of  the  container  must  be 
low.     This  eliminates  gaseous  fuels. 

The  liquid  fuels  which  meet  the  above  conditions  best  are: 
(a)  certain  hydrocarbons,  which  form  the  constituents  of  gasoline 
and  of  certain  coal  tar  products,  and  (b)  the  alcohols.  The 
hydrocarbons  under  consideration  may  be  divided  into  two  main 
groups,  saturated  and  unsaturated.  The  latter  term  is  here 
applied  to  the  behavior  and  not  to  the  composition  of  the  sub- 
stance. The  saturated  hydrocarbons  are  again  subdivided  into 
the  aliphatic  or  acyclic  group,  and  into  the  aromatic  or  cyclic 
group.  The  hydrocarbons  all  form  series  in  which  the  members 
differ  from  each  other  by  the  addition  of  CH2.  The  members  of 
the  groups  of  most  importance  are  listed  in  Table  9,  together  with 

212 


FUELS  AND  EXPLOSIVE  MIXTURES 


213 


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Pentane 
Hexane 
Heptane 
Octane 

Isopentane 
Isohexane 
Isoheptane 
Isooctane 

Benzene 
Toluene 
Xylene 

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214  THE  AIRPLANE  ENGINE 

the  boiling  points  (temperatures  of  vaporization  at  atmospheric 
pressure),  densities  as  compared  with  that  of  water  and  latent 
heats  of  evaporation. 

Certain  properties  of  these  compounds  are  given  in  the  follow- 
ing tables,  in  which  there  are  also  included,  for  convenience,  the 
properties  of  air,  and  its  constituents,  and  of  the  products  of 
combustion  of  the  fuel  elements.  In  Table  10  data  are  given  for 
the  gaseous  state  only.  Column  3  gives  the  density  of  each 
substance  compared  with  air  at  the  same  pressure  and  tempera- 
ture; column  4  gives  the  weight  in  pounds  per  cubic  foot  of  the 
substance.  Some  of  the  quantities  there  given  are  fictitious  in 
that  the  substance  is  not  gaseous  at  32°F.  and  760  mm.,  but  the 
quantity  is  of  value  in  permitting  the  easy  determination  of 
specific  weight  at  those  temperatures  and  pressures  at  which  it  is 
gaseous.  Column  5  is  the  reciprocal  of  column  4.  The  quantity 
R  is  the  constant  in  the  gas  equation  pv  =  wRT. 

In  Table  11  are  given  combustion  data  for  the  fuel  constituents. 
Column  4  gives  the  volume  of  air  necessary  to  burn  one  volume  of 
the  gaseous  fuel,  both  being  at  the  same  pressure  and  tempera- 
ture; the  products  of  complete  combustion  are  in  all  cases  C02, 
H20  and  N2  and  their  volumes  are  given  in  columns  5,  6  and  7. 
The  mixture  usually  experiences  a  change  in  volume  as  a  result 
of  the  chemical  changes  resulting  from  combustion,  entirely 
independent  of  the  change  in  pressure  and  temperature;  this 
change  in  volume  is  given  in  column  8.  Columns  9  to  12  give  the 
weight  of  air  required  for  combustion  of  1  Ib.  of  fuel  and  the 
weights  of  each  of  the  resulting  products. 

Table  12  gives  the  heat  of  combustion  of  each  of  the  substances 
listed.  There  is  also  given  the  heat  of  combustion  per  pound 
of  explosive  mixture,  and  the  heat  of  combustion  per  cubic  foot 
of  explosive  mixture  measured  at  60°F.  and  standard  atmospheric 
pressure,  the  mixture  being  assumed  to  contain  only  that  amount 
of  air  which  is  chemically  necessary.  It  will  be  noted  that  two 
values  of  heat  of  combustion  are  given  under  each  head,  a  higher 
and  a  lower  heat  value,  and  that  these  are  different  for  all  the 
fuels  which  contain  hydrogen.  The  higher  heat  value  is  the 
total  heat  liberated  by  combustion,  or  the  heat  which  would 
be  given  up  by  the  mixture  when  burned  in  a  closed  vessel 
and  cooled  to  its  initial  temperature.  Whenever  there  is 
hydrogen  in  the  fuel,  water  is  formed  by  combustion  and  part 
of  the  heat  of  combustion  is  absorbed  in  vaporizing  it.  If, 


FUELS  AND  EXPLOSIVE  MIXTURES 


215 


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216 


THE  AIRPLANE  ENGINE 
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KNO-KNOOO--IOO— lOO 


FUELS  AND  EXPLOSIVE  MIXTURES 


217 


TABLE  12. — HEATS  OP  COMBUSTION 


Substance 

Chemical 
formula 

B.t.u.  per  pound 

B.t.u.  per  cubic 
foot  of  theoreti- 
cal mixture  at 
60°F.  and  760 
mm. 

B.t.u.  per 
pound    of   theo- 
retical mixture 

High 

Low 

High 

Low 

High 

Low 

H2 
CO 

CH4 
C2H8 
CaH8 
C4Hio 
C6Hi2 
CsHu 
C7Hi6 
CsHis 
C»H2o 
CaHe 
C7H8 
C«Hu 
C2H4 
C3H6 
C4H8 
C2H* 
C3H4 
CioH8 
CH4O 
C2H6O 

62,100 
4,380 
23,850 
22,230 
21,600 
21,240 
21,140 
20,800 
20,600 
20,400 
20,380 
18,070 
18,250 

52,920 
4,380 
21,670 
20,500 
20,055 
19,780 
19,600 
19,380 
19,200 
19,020 
19,015 
17,400 
17,490 
18,900 
20,420 
20,150 
19,700 
21,020 
20,325 
16,860 
8,460 
11,650 

95.6 
94.6 
95.8 
99.7 
101.2 
101.7 
103.0 
102.5 
102.0 
101.4 
101.7 
101.0 
100.8 

104.9 
104.8 
104.4 
115.0 
112.1 
101.1 
99.5 
105.2 

81.5 
94.6 
87.0 
91.8 
94.0 
94.6 
95.6 
95.4 
95.0 
94.5 
95.0 
97.2 
96.5 
95.5 
99.2 
99.3 
98.5 
112.0 
107.0 
98.0 
88.2 
94.3 

,760 
,265 
,310 
,300 
,297 
,285 
,297 
,284 
,275 
,266 
,266 
,184 
,261 

,371 
,356 
,325 
,515 
,434 
,250 
,281 
,303 

,500 
,265 
,190 
,200 
,205 
,203 
,203 
,197 
,188 
,176 
,184 
,140 
,208 

,295 
,280 
,254 
,475 
,375 
,210 
,133 
1,167 

Carbon-monoxide  
Methane 

Ethane  

Propane  

Butane  
Pentane  
Hexane  
Heptane  

Toluene 

Cyclo-hexane  .  .  . 

Ethylene  .  .  .... 

21,600 
21,330 
20,880 
21,600 
21,200 
17,410 
9,550 
13,000 

Propylene  

Butylene  
Acetylene  
Allylene  
Naphthalene  
Methyl  alcohol  
Ethyl  alcohol 

as  is  usual  in  gas  engines,  the  gases  escape  at  so  high  a  tem- 
perature that  no  water  vapor  is  condensed  in  the  cylinder, 
the  latent  heat  of  vaporization  of  the  water  is  not  available  for 
raising  the  temperature  of  the  products  of  combustion,  or  for 
doing  work.  The  useful  heat  of  combustion  is  consequently  the 
total  heat  less  the  heat  absorbed  in  vaporizing  the  H2O  formed 
by  the  combustion.  The  heat  of  vaporization  depends  on  a 
number  of  factors.  A  value  of  950  B.t.u.  per  pound  may  be 
assumed  and  if  this  is  multiplied  by  the  number  of  pounds  of 
water  formed  per  pound  of  fuel  burned  it  will  give  (with  an 
approximation  adequate  for  ordinary  purposes)  the  unavailable 
heat.  The  lower  heat  value  is  the  total  heat  minus  the  unavail- 
able heat  and  is  the  value  commonly  used  by  engineers  in  dealing 
with  gas  engine  problems. 

Gasoline. — The  gasolines  at  present  on  the  market  are  of  three 
different  types:1 

1  DEAN,  Motor  Gasoline,  Technical  Paper  166,  U.  S.  Bureau  of  Mines. 


218  THE  AIRPLANE  ENGINE 

1.  " Straight"  refinery  gasoline. 

2.  Blended  casing-head  gasoline. 

3.  Cracked  and  blended  gasoline. 

"Straight"  refinery  gasoline  is  produced  by  distillation. 
Crude  petroleum  is  first  distilled  from  a  fire  still,  and  the  con- 
densed product  is  collected  until  it  reaches  some  predetermined 
density.  This  so-called  crude  naphtha  or  benzine  is  then  acid- 
refined  and  steam-distilled.  Several  products  of  different  ranges 
of  volatility  may  be  produced  or  the  steam  distillation  may  simply 
separate  the  product  from  the  less  volatile  " bottoms."  Straight 
refinery  gasolines  consist  mainly  of  aliphatic  hydrocarbons  (see 
Table  9)  and  are  generally  characterized  by  a  low  content  of 
unsaturated  and  aromatic  hydrocarbons  and  by  a  distillation 
range  free  from  marked  irregularities. 

Blended  casing-head  gasolines  are  of  recent  development. 
Casing-head  gasoline  is  obtained,  by  compression  or  absorption, 
from  natural  gas  and  is  too  volatile  for  general  use.  Before 
marketing,  it  is  generally  blended  with  enough  heavy  naphtha  to 
produce  a  mixture  that  can  be  handled  safely.  As  a  result  of  this 
blending,  the  volatility  range  is  usually~characterized  by  a  con- 
siderable percentage  of  constituents  of  both  low  and  high  boiling 
points  and  a  lack  of  intermediate  constituents.  Skilful  blending 
may  change  this  characteristic. 

In  chemical  properties  the  blended  casing-head  gasoline  seems 
to  be  identical  with  a  straight  refinery  product  of  the  same 
distillation  range. 

Cracked  or  synthetic  gasoline  is  also  a  recent  development. 
An  oil  consisting  mainly  of  heavier  hydrocarbons  is  subjected  to 
high  temperature  and  pressure  and  is  thereby  broken  down  or 
" cracked"  into  lighter  constituents.  This  cracked  gasoline 
is  generally  marketed  in  the  form  of  blends  with  refinery  and 
casing-head  gasoline.  Cracked  gasolines  differ  chemically  from 
straight-refinery  and  casing-head  gasolines  in  having  a  larger 
amount  of  unsaturated  and  aromatic  hydrocarbons.  The  heat 
of  combustion  of  the  aromatic  compounds  averages  about  15  per 
cent  less  than  that  of  the  acyclic  compounds,  but,  as  shown  on 
page  238,  the  presence  in  moderate  amounts  of  certain  aromatic 
compounds  may  improve  the  thermal  efficiency  of  the  engine 
enough  to  offset  any  disadvantage,  for  airplane  use,  of  a  lower 
heat  of  combustion. 


FUELS  AND  EXPLOSIVE  MTXTRES 


219 


Specifications. — In  the  past,  gasolines  have  usually  been 
described  and  bought  on  a  gravity  specification.  So  long  as  the 
gasolines  in  the  market  were  straight  refinery  products  such  a 
specification  was  reasonably  satisfactory,  but,  with  the  develop- 
ment of  blended  casing-head  gasolines,  it  has  become  impossible 
to  determine  the  volatility  of  a  gasoline  by  density  measurements. 
A  given  density  may  represent  either  a  narrow  "cut"  consist- 
ing of  a  product  which  evaporates  with  a  very  narrow  range  of 
temperature,  or  a  mixture  of  a  volatile  low-density  product  with  a 
product  of  high  density  and  low  volatility.  The  former  fuel 
might  be  admirable  for  airplanes  while  the  latter  might  be  quiet 
unsuitable. 

Specific  gravity  is  best  expressed  as  the  ratio  of  the  density 
of  the  fuel  to  that  of  water,  both  at  60°F.  The  trade  practice 
has  been  to  use  the  Baume*  scale  of  density.  This  arbitrary 
scale  has  nothing  to  recommend  it,  and  suffers  the  disadvantage 


TABLE  13. — SPECIFIC  GRAVITIES  AT 


60g 
60C 


F.  CORRESPONDING  TO  DEGREES 


BAUM£  FOR  LIQUIDS  LIGHTER  THAN  WATER 


1 

'? 

1 

| 

1 

1 

1 

I 

1 

* 

1 

! 

I 

1 

B 

1 

I 

• 

i 

1 

1 

* 

1 

1 

I 

1 

1 

& 

8.FH 

$ 

P 

P 

I 

1 

t 

P 

I 

I 

Q 

% 

25 

0.9032 

40 

0.8235 

55 

0.7568 

70 

0.7000- 

85 

0.6512 

26 

0.8974 

41 

0.8187 

56 

0.7527 

71 

0.6965 

86 

0.6482 

27 

0.8917 

42 

0.8140 

57 

0.7487 

72 

0.6931 

87 

0.6452 

28 

0.8861 

43 

0.8092 

58 

0.7447 

73 

0.6897 

88 

0.6422 

29 

0.8805 

44 

0.8046 

59 

0.7407 

74 

0.6863 

89 

0.6393 

30 

0.8750 

45 

0.8000 

60 

0.7368 

75 

0.6829 

90 

0.6364 

31 

0.8696 

46 

0.7955 

61 

0.7330 

76 

0.6796 

91 

0.6335 

32 

0.8642 

47 

0.7910 

62 

0.7292 

77 

0.6763 

92 

0.6306 

33 

0.8589 

48 

0.7865 

63 

0.7254 

78 

0.6731 

93 

0.6278 

34 

0.8537 

49 

0.7821 

64 

0.7216 

79 

0.6699 

94 

0.6250 

35 

0.8485 

50 

0.7778 

65 

0.7179 

80 

0.6667 

95 

0.6222 

36 

0.8434 

51 

0.7735 

66 

0.7143 

81 

0.6635 

96 

0.6195 

37 

0.8383 

52 

0.7692 

67 

0.7107 

82 

0.6604 

97 

0.6167 

38 

0.8333 

53 

0.7650 

68 

0.7071 

83 

0.6573 

98 

0.6140 

39 

0.8284 

54 

0.7609 

69 

0.7035 

84 

0.6542 

99 

0.6114 

100 

0  .  6087 

220 


THE  AIRPLANE  ENGINE 


that  the  greater  the  density  the  lower  is  the  number  of  "degrees" 
on  the  Baume*  scale.  For  liquids  lighter  than  water,  the  relation 
between  the  specific  gravity  and  Baume  scale,  B,  is  given  by 
the  expression 

Specific  gravity  = 


Numerical  values  are  given  in  Table  13.  Commercial  gasolines 
range  from  about  55  to  75°Be\  (Sp.  gr.  0.758  to  0.684).  The 
Eastern  gasolines  are  lightest  (60  to  75°Be\)  and  the  California 
gasolines  heaviest  (57  to  63°Be\). 


Cork 

Thermometer 

Cooling  Jnough 
Cork  /  Containing  Cracked  Ice  and  Wafer 


FIG.  164. — Distillation  apparatus. 

Volatility. — Volatility  is  the  basic  property  that  determines  the 
grade  and  usefulness  of  a  gasoline.  The  presence  of  low-boiling 
constituents  is  desirable  to  permit  easy  starting  of  a  cold  motor 
but  may  result  in  excessive  evaporation  losses  in  the  commercial 
handling  of  the  fuel. 

The  volatility  is  determined  by  distillation  (Fig.  164).  A 
100-gram  sample  of  the  fuel  is  heated  slowly  while  the  vapor  given 
off  is  condensed  and  collected.  The  first  drop  of  gasoline  should 
fall  from  the  end  of  the  condenser  tube  in  5  to  10  min.  The 


FUELS  AND  EXPLOSIVE  MIXTURES 


221 


rate    of    evaporation    is    kept    about    4    c.c.   per  minute.     A 

thermometer  indicates  the  temperature  of  the  vapor  above  the 

fuel.     Readings    of 

this    temperature    are 

taken  when   the   first 

drop  of  distillate  falls 

(initial  point)  and  as 

each    10   per  cent  or 

other     selected     p  e  r- 

centage   has  distilled, 

until    at    the    end    a 

dry  point  is  reached. 

The  observations   are 

usually  plotted  as  in 

Fig.    165   and  give  a 

record  of  the  volatility 

of  the  fuel. 

Specifications      for 
aviation  gasoline  have 
been  prepared  by  the  U.  S.  Committee  on  Standardization  of 
Petroleum  Specifications1  and  are  as  follows: 

GASOLINE  DISTILLATION  TEST  SPECIFICATIONS 


\\J\J 
90 
30 

-.70 

0> 

o 

10, 

•i 

FIG.  1 

x^ 

^ 

J 

^ 

^ 

/ 

'' 

4 

/ 

'''$ 

?x 

/ 

/ 

VV'/ 

0| 

'/ 

/ 

// 

1 

1 

tj- 

/  <b- 

t 

/; 

3 

/ 

/ 

t 

4 

P 

A 

4 

—  Eastern  Sfa/ght  -Rtf  'inert, 
Gasoline 
—  Blended  Casi'ng-head  Oaso/ine 

r 

/-' 

* 

/  - 

x 

0             75              100            125             150            20 
Temperature  ,  Deg.Cent 

65.  —  Distillation  curves  for  straight  refiner 
and  casing-head  gasoline. 

Grade 

Aviation 
gasoline, 
domestic  grade 

Aviation 
gasoline, 
fighting  grade 

Thermometer  reading  range  when  5  per 
cent  is  recovered  in  receiver     

122  to  167°F. 

122  to  149°F. 

Thermometer  reading  when  50  per  cent  is 
recovered  in  receiver,  not  more  than  .... 
Thermometer  reading  when  90  per  cent  is 
recovered   in  receiver,  not  more  than 
Thermometer  reading  when  96  per  cent  is 
recovered  in  receiver,  not  more  than  .... 
End-point  shall  not  be  higher  than  

221°F. 
311°F. 

347°F. 
374°F. 

203°F. 
257°F. 

302°F. 
329°F. 

Distillate  recovered  in  the  receiver  from 
the  distillation  at  least  

96  per  cent 

96  per  cent 

When  the  residue  is  cooled  and  added  to 
the  distillate  in  the  receiver  the  distilla- 
tion loss  shall  not  exceed 

2  per  cent 

2  per  cent 

1  Bureau  of  Mines,  Bulletin  No.  5,  effective  Dec.  29,  1920. 


222  THE  AIRPLANE  ENGINE 

In  addition  the  specifications  require  for  both  grades  of 
aviation  gasoline  the  following  properties: 

Color:  Water  white. 

Doctor  test :  Negative. 

Corrosion  test:  100  c.c.  of  the  gasoline  shall  cause  no  gray  or  black 
corrosion  and  no  weighable  amount  of  deposit  when  evaporated  in  a  polished 
copper  dish. 

Unsaturated  hydrocarbons:  maximum  proportion  of  the  gasoline  soluble 
in  concentrated  sulphuric  acid,  2  per  cent. 

Acid  heat  test:  the  gasoline  shall  not  increase  in  temperature  more  than 
10°F. 

Acidity:  the  residue  after  distillation  shall  not  show  an  acid  reaction. 

The  gasoline  shall  be  free  from  undissolved  water  and  suspended  matter. 

The  Doctor  Test  is  made  by  shaking  two  volumes  of  gasoline 
with  one  volume  of  "doctor"  solution  (sodium  plumbite)  in  a 
test  tube,  shaking  for  15  sec.,  adding  a  pinch  of  flowers  of  sulphur, 
shaking  again  and  allowing  to  settle.  If  the  liquid  remains 
unchanged  in  color  and  the  sulphur  remains  bright  or  only  slightly 
discolored,  the  test  is  negative  and  the  gasoline  is  " sweet." 

The  Acid  Heat  Test  is  made  by  adding  30  c.c.  of  66°  commercial 
sulphuric  acid  to  150  c.c.  of  gasoline,  both  being  at  room 
temperature.  After  mixing,  shake  for  2  min.  and  observe  the 
rise  in  temperature. 

Volatility  curves  for  three  straight  refinery  gasolines  and  for 
three  blended  casing-head  gasolines,  of  approximately  the  same 
densities,  are  given  in  Fig.  165.  The  casing-head  gasolines  are  seen 
to  have  larger  percentages  distilled  below  50°C.,  but  have  longer 
distillation  ranges.  This  results  in  a  fairly  uniform  slope  of  the 
distillation  curve.  The  large  percentage  unevaporated  at  150°C. 
shows  that  the. fuel  is  a  blend  or  mixture  of  heavier  and  lighter 
fuels. 

It  should  be  noted  further  that  the  "cut"  or  fraction  distilling 
off  at  any  given  temperature  will  be  different  from  different 
gasolines.  This  is  demonstrated  in  Fig.  166,  which  shows  that  the 
100°C.  cut  may  vary  in  specific  gravity  from  0.710  to  0.733  and  the 
150°C.  cut  from  0.748  to  0.780.  In  other  words,  volatility  alone 
is  not  sufficient  to  characterize  a  gasoline. 

From  the  volatility  curves  for  straight  refinery  gasoline,  Fig. 
165,  it  will  be  seen  that  the  average  temperature  of  evaporation 
(boiling  temperature)  from  a  high-grade  gasoline  is  100°C.  From 
the  specific  gravity  curves  it  appears  that  the  average  density  at 


FUELS  AND  EXPLOSIVE  MIXTURES 


223 


100°C.  is  almost  0.7.  As  the  constituents  are  mainly  aliphatic 
hydrocarbons  it  is  safe  to  assume  that  a  high-grade  gasoline  con- 
sists principally  of  hexane  (C6Hi4)  and  heptane  (C7Hi6),  whose 
boiling  points  are  69  and  98.4°C.  and  densities  0.676  and  0.7,  re- 
spectively (see  Table  9). 
Calorific  Value.— The 
calorific  value  of  com- 
mercial gasoline  varies 
very  slightly  with  type 
of  fuel,  field  of  origin,  or 
density.  Exhaustive 
tests  by  the  U.  S.  Bureau 
of  Mines  show  only  1.5 
per  cent  difference  be- 
tween the  highest  and 
lowest  values  for  a  range 
of  density  from  0.687  to 
0.745  (73.8  to  57.9°Be\), 
the  samples  investigated 
including  all  the  com- 
mercial types.  The 
average  high  heat  value 
is  20,200  B.t.u.  per 
pound.  It  should  be 
noted,  however,  that 
gasoline  is  sold  by  the 
gallon  and  that  there  is 
considerable  difference 


—Eastern  Straight Jfcfinery 

6asof/ne. 
Mid-Continent  Stmight 

Refinery  Gasoline.  

CaliforniaStraigfrt  Refinery 
Oaso/irre. 


tmBlended  Casing-head  (Eastern) 
"Gasoline. 

Cracked  (Mid-  Continent) 

Gasoline. 


15          100         125          150 
Temperature,  DecjCent 

FIG.  166. — Density  of  "cuts"  from  various  gas- 
olines. 


on  that  basis;  the  calorific  value  ranges  from  124,000  B.t.u.  per 
gallon  for  sp.  gr.  0.745,  to  116,500  B.t.u.  for  sp.  gr.  0.687,  a  dif- 
ference of  over  7  per  cent  in  favor  of  the  heavier  fuel.  This 
difference  is  not  important  in  airplane  use,  since  the  weight  of 
fuel  that  has  to  be  carried  is  the  important  factor,  and  not  its 
volume;  in  automobile  use  it  may  more  than  offset  the  disad- 
vantages resulting  from  the  use  of  a  less  volatile  fuel. 

Benzene  or  benzol  (C6H6)  is  a  fuel  which  has  been  used  con- 
siderably in  airplanes,  though  generally  mixed  with  gasoline.  It 
is  obtained  chiefly  from  by-product  coke-ovens. 

Commercial  90  per  cent  benzol  has  a  specific  gravity  of  about 
0.88.  Its  distillation  curve  should  show  an  initial  point  not 
lower  than  74°C.,  90  per  cent  at  or  below  86°C.,  95  per  cent  at  or 


224 


THE  AIRPLANE  ENGINE 


below  95°C.,  and  end  point  not  above  150°C.  With  a  calorific 
value  of  18,000  B.t.u.  per  pound  the  heating  value  per  gallon  is 
132,000  B.t.u.,  or  considerably  higher  than  that  of  gasoline. 

When  mixed  with  gasoline  there  is  no  change  in  total  volume. 
The  distillation  curve  for  such  a  mixture,  containing  20  per  cent 
of  benzol  and  80  per  cent  high-grade  gasoline,  is  shown  in  Fig. 
167,  together  with  the  distillation  curves  of  the  benzol  and  the 
gasoline. 


First  10 
Drop 


20     30 


80      90 


Dry 


40       50      60      70 

Percentage  Distilled  Point 

FIG.   167. — Distillation  curve  of  benzol-gasoline  mixture. 


It  will  be  observed  that  the  effect  of  the  addition  of  benzol  is 
to  increase  the  volatility  of  the  mixture;  with  30  to  50  per  cent 
distilled  the  volatility  is  greater  than  that  of  either  of  the  con- 
stituents, after  which  it  becomes  intermediate  to  the  volatility  of 
the  constituents.  This  improvement  in  volatility  has  been 
found  to  be  distinctly  advantageous  in  increasing  engine  power 
at  high  altitudes.  The  lower  heat  of  combustion  of  the  mixture 
results  in  the  consumption  of  a  greater  weight  of  fuel  per  brake 
horse-power  hour  than  with  straight  gasoline. 

Alcohol  has  been  used  mixed  with  gasoline  or  benzol  or  both,  as 
an  airplane  fuel.  Methyl  alcohol,  CH40  (wood  alcohol)  has  a 
boiling  point  of  65°C.  and  sp.  gr.  0.81,  heats  of  combustion,  high 
9,550,  low  8,460,  B.t.u.  Ethyl  alcohol,  C2H60  (grain  alcohol) 
has  a  boiling  point  of  78°C.,  density  0.79,  heats  of  combustion, 
high  13,000,  low  11,650  B.t.u.  Commercial  alcohol,  either  pure 
or  denatured,  contains  water  (e.g.,  10  per  cent  by  volume  in  90 


FUELS  AND  EXPLOSIVE  MIXTURES  225 

per  cent  alcohol)  and  has  a  higher  boiling  point  than  pure  alcohol. 
The  effect  of  the  addition  of  alcohol  to  gasoline  is  to  improve  the 
volatility.  The  calorific  value  of  alcohol  is  so  low  compared  with 
gasoline  that  its  use  inevitably  increases  the  weight  of  fuel 
burned  per  unit  of  power  developed.  It  does  not,  however, 
diminish  the  power  developed,  because  the  heat  of  combustion 
per  unit  volume  of  explosive  mixture  (see  Table  12)  is  about  the 
same  as  for  gasoline;  it  may  even  increase  the  power  output 
slightly.  Its  use  also  permits  an  increase  in  the  permissible 
compression  ratio  for  the  engine  and  thereby  improves  the 
thermal  efficiency. 

Hydrogen  has  the  highest  calorific  value  of  any  of  the  fuels,  per 
pound,  but  not  per  cubic  foot  of  explosive  mixture  (Table  12). 
Apart  from  its  high  cost,  it  is  objectionable  on  account  of  the 
great  violence  of  the  explosion  when  mixed  with  the  proper 
amount  of  air.  It  cannot  be  carried  in  airplanes  unless  com- 
pressed to  very  high  pressure  in  steel  tanks,  which  makes  the 
fuel  system  too  heavy,  or  in  the  liquid  form,  which  increases 
greatly  the  cost  of  the  fuel.  Liquid  hydrogen  has  a  temperature 
below  —  400°F.  and  cannot  be  kept  from  evaporating  rapidly 
without  the  very  best  of  heat  insulation;  no  sufficiently  robust 
container  with  adequate  heat-insulating  qualities  has  been  devised 
as  yet.  Hydrogen  gas  has  been  used  for  starting  cold  engines. 

It  is  often  necessary  to  waste  some  of  the  hydrogen  contained 
in  a  dirigible  balloon.  Attempts  to  burn  the  hydrogen  alone  in 
the  engine  have  shown  that  only  about  one-third  of  the  maximum 
horse  power  of  the  engine  could  be  developed  without  serious 
detonations.  By  mixing  hydrogen  with  the  gasoline  it  is  possible 
to  develop  the  maximum  power  without  trouble  and  thereby  to 
save  gasoline;  at  the  higher  powers  only  a  small  quantity  of 
hydrogen  can  be  burned. 

Acetylene,  (C2H2),  like  Hydrogen,  gives  explosions  of  great 
violence  in  the. cylinder.  Its  heat  of  combustion  per  cubic  foot 
of  explosive  mixture  is  highest  of  all  the  fuels  given  in 
Table  12.  It  may  be  stored  either  in  the  gaseous  or  liquid 
forms,  but  with  the  same  objections  (though  to  a  less  degree) 
as  hydrogen.  It  can  be  generated  by  adding  water  to  calcium 
carbide,  leaving  a  residue  of  slaked  lime.  As  the  residue  is 
considerably  heavier  than  the  acetylene,  the  total  weight  of  the 
fuel  system  becomes  excessive,  if  it  is  attempted  to  generate  the 
acetylene  in  an  airplane. 

15 


226  THE  AIRPLANE  ENGINE 

Ether  has,  as  its  principal  advantage,  the  fact  that  its  boiling 
point  (35°C.)  is  lower  than  that  of  any  of  the  other  available  fuels 
which  are  liquid  at  ordinary  temperatures.  This  gives  it  a  special 
value  in  starting  a  cold  engine.  Its  use  has  been  restricted  to 
that  purpose.  The  heat  of  combustion  is  rather  low. 


EXPLOSIVE  MIXTURES 

Properties  of  Vapors. — Every  liquid  gives  off  vapor  continu- 
ously until  the  pressure  exerted  by  that  vapor  at  the  surface  of  the 
liquid  reaches  a  limiting  value,  which  depends,  for  any  given 
liquid,  on  its  temperature  only.  The  vapor  liberated  is  always 
at  the  temperature  of  the  liquid  and  it  is  said  to  be  " saturated" 
when  it  is  at  the  limiting  pressure.  The  relation  between  the 
pressure  and  temperature  of  a  saturated  vapor  is  determinable 
only  by  experiment. 

The  pressure  exerted  by  a  vapor  will,  in  time,  reach  the  satura- 
tion pressure  if  the  liquid  is  contained  in  a  vessel  of  moderate 
dimensions;  the  presence,  above  the  liquid,  of  gases  or  other 
vapors  which  are  inert  to  the  vapor  under  consideration  and  are 
at  the  same  temperature  will  not  affect  the  saturation  pressure. 
The  total  pressure  in  the  vessel  (assuming  no  change  of  tempera- 
ture) will  be  the  sum  (1)  of  the  pressures  of  the  gases  and  vapors 
already  there  and  (2)  of  the  saturation  pressure  of  the  liquid. 

If  the  containing  vessel  is  very  large  or  if  the  time  available 
is  too  short,  or  if  the  weight  of  liquid  put  into  the  vessel  is  less 
than  the  weight  of  saturated  vapor  necessary  to  fill  the  vessel, 
the  vapor  will  have  a  pressure  less  than  the  saturated  pressure 
and  will  be  in  the  condition  known  as  "superheated."  Suppose 
the  temperature  of  the  superheated  vapor  to  be  T.  If  this 
vapor  is  cooled  at  constant  pressure,  with  consequent  diminu- 
tion in  volume,  a  temperature,  T0,  will  eventually  be  reached 
at  which  the  vapor  is  saturated.  The  cooling  process  is  similar 
to  that  used  for  determining  the  dew-point  of  air;  the  dew-point 
is  the  saturation  temperature.  The  vapor  is  said  to  be  super- 
heated T-T0  degrees.  All  unsaturated  vapors  are  superheated. 
When  superheated  they  may  be  considered  to  behave  like  per- 
fect gases. 

A  saturated  vapor  cannot  exist,  as  such,  at  a  pressure  greater 
than  the  saturation  pressure.  If  a  saturated  vapor  is  cooled  at 


FUELS  AND  EXPLOSIVE  MIXTURES  227 

constant  volume,  thereby  lowering  the  saturation  pressure, 
some  of  the  vapor  will  condense.  If  a  saturated  vapor  is  com- 
pressed, keeping  the  temperature  constant,  condensation  will 
take  place;  if,  on  the  other  hand,  it  is  expanded  at  constant 
temperature  it  will  become  unsaturated  (superheated)  unless 
liquid  is  present  to  supply  more  vapor  by  evaporation.  The 
presence  of  other  inert  vapors  will  not  affect  these  phenomena. 
If  air  is  passed  through  or  over  a  liquid  (as  in  certain  obsolete 
types  of  carburetor),  and  if  the  contact  is  sufficiently  intimate 
and  prolonged,  the  air  will  leave  carrying  with  it  the  saturated 
vapor  of  the  liquid.  If  a  liquid  is  injected  into  a  current  of  air 
(as  in  modern  carburetors)  and  if  the  contact  is  sufficiently 
intimate  and  prolonged  and  if,  furthermore,  the  weight  of  liquid 
present  is  sufficient  for  that  purpose,  the  air  will  carry  its  own 
volume  of  the  saturated  vapor  of  the  liquid.  If  more  liquid  is 
injected  than  is  necessary  for  this  purpose  the  excess  liquid  will 
remain  in  the  liquid  state.  In  any  case,  the  total  pressure  of  the 
carbureted  air  is  the  sum  of  the  partial  pressures  of  the  air  and  of 
the  vapor.  If  the  pressure  of  the  carbureted  mixture  is  p,  and 
the  pressure  of  the  vapor  is  pv,  and  of  the  air  in  the  carbureted 
mixture  is  pa,  then 

p  =  Pa  +  Pv 

The  relation  between  the  saturation  pressures  and  temperatures 
of  the  liquid  fuels  which  are  of  importance  in  airplane  engines  is 
given  in  Fig.  168.  Table  14  gives  their  values  for  certain  selected 
temperatures. 

The  specific  volumes  (volumes  of  1  Ib.)  of  the  saturated  vapors 
are  calculated  on  the  assumption  that  they  are  perfect  gases. 
This  assumption  is  fairly  satisfactory  for  the  low  vapor  pressures 
which  alone  are  of  interest  in  engine  mixtures.  Taking,  for 
example,  heptane  (C7Hi6)  at  60°F.,  the  molecular  weight  is  7 
X  12  +  16  =  100.  The  gas  constant  R  is  inversely  as  the  mo- 
lecular weight  of  the  gas;  taking  R  for  oxygen  as  48.25,  R  for  hep- 

32 
tane  is  T^  X  48.25  =  15.45,  and  the  specific  volume  at  60°F. 

1UU 

and  at  the  saturation  pressure  of  0.54  Ib.  per  square  inch  is 
RT       15.45  X  520 


0.54  X  144 


=  10B 


228 


THE  AIRPLANE  ENGINE 


The  weight  of  air  required  for  combustion  is  obtained  from  the 
chemical  equation, 

C7Hi6  +  1102  =  7CO2  +  8H2O 

The  relative  weights  of  heptane  and  oxygen  are  100  and  11  X  32. 
As  the  oxygen  content  of  air  is  23.4  per  cent  by  weight,  the  air 


11  X32 
100 


X 


required  for  the   combustion  of  1  Ib.  of  heptane  is 

100 
^  =  15.1  Ib.     The  volume  of  this  air  at  60°F.  and  14.7  Ib. 


°30    40     50     60     70      80    90     100    110    120   '130    140    150     160    170    180 

Temperature,  Decj.  Fahr 
FIG.   168. — Vapor  pressures  of  various  liquid  fuels. 

wRT 
per  square  inch  pressure  is  given  by  the  equation  v  =  — — 

15.1  X  53.4  X  520  =197cufl 


4x44 

It  is  desirable  that  the  fuel  entering  the  inlet  manifold  should 
be  entirely  in  the  vapor  form,  either  superheated  or  just  saturated. 
This  condition  is  necessary  to  obtain  a  homogeneous  mixture  and 
an  equal  distribution  of  the  fuel  to  all  the  cylinders.  The  possi- 
bility of  obtaining  this  condition  may  be  determined  by  continu- 
ing the  preceding  calculation.  If  the  air  is  saturated  with 
heptane  vapor  at  60°F.  its  partial  pressure,  pa,  will  be  14.7  —  0.54 
=  14.16  Ib.  per  square  inch  and  the  volume  of  the  air  at  this 


FUELS  AND  EXPLOSIVE  MIXTURES 


229 


Pi 

ni 


11 


«  "3  p,^ 

i«i' 

1=2 

M    m    tt 


A 


a 


O    02^ 

II* 

CL    S     (n    4^ 

rr\      «      ii      rrt 


'8 


a-s 


oooooooo 


SO) 
• 


rHOOCOCO 
OOOt^OO 


OOOOOOOOi-lWr-l 


lOi-HOOOOOOO 


cocot^t>.t>.t^osoooo 
doooddddo' 


2   S 


.z  *        M 

OOOOOOOO  0 


i  i  ;1 


230  THE  AIRPLANE  ENGINE 

14.7 
reduced  pressure  will  be  TT~TB  X  197  =  202  cu.  ft.     This  quantity, 

and  similar  quantities  for  other  air  temperatures  and  other  fuels, 
are  given  in  Table  14.  The  vapor  coexists  in  the  same  space  with 
the  air  (202  cu.  ft.)  and  as  this  volume  is  greater  than  the  volume 
which  the  saturated  vapor  of  heptane  occupies  (103  cu.  ft.)  the 
vapor  cannot  be  saturated  in  a  chemically  correct  mixture;  the 
vapor  will  be  superheated.  At  the  temperature  of  40°F.  the  air 
volume  is  seen  to  be  194  cu.  ft.  and  that  of  the  saturated  vapor  of 
heptane  185  cu.  ft.;  as  these  are  approximately  equal  the  vapor 
will  be  practically  saturated.  At  temperatures  lower  than  40°F. 
the  volume  of  the  air  will  be  less  than  the  volume  of  the  saturated 
vapor  and  in  that  case  part  of  the  fuel  will  necessarily  be  in  the 
liquid  form.  An  excess  of  air  above  that  chemically  necessary 
will  lower  the  temperature  at  which  liquid  must  begin  to  appear; 
an  excess  of  fuel  will  raise  that  temperature. 

With  a  less  volatile  fuel  such  as  octane  it  will  be  seen  by 
inspection  of  the  table  that  a  higher  temperature  (a  little  under 
80°F.)  will  be  necessary  if  the  fuel  is  to  be  in  the  vapor  form. 
With  benzol  the  temperature  is  well  below  40°;  with  methyl  and 
ethyl  alcohol  between  70  and  80°F. 

It  should  be  noted  that  the  temperatures  of  the  table  are  the 
'temperatures  after  the  vapor  is  formed.  In  the  carburetor,  the 
latent  heat  of  vaporization  of  the  fuel  is  taken  from  the  air  and 
the  liquid  fuel,  with  the  result  that  the  temperature  of  the  mixture 
falls  below  the  temperature  of  the  entering  air  and  fuel,  unless 
heat,  equal  to  the  latent  heat,  is  supplied  from  the  jacket  water  or 
exhaust  gases.  If  the  latent  heat  of  the  fuel  is  135  B.t.u.,  the 
specific  heat  of  the  liquid  0.45  and  the  specific  heat  of  air  at 
constant  pressure  0.241,  the  fall  in  temperature  AT7  for  a  mixture 
of  1  Ib.  of  fuel  with  15.1  Ib.  of  air  is  given  by  the  equation 

135  =  A7X0.45  +  15.1  X  0.241) 
or  AT7  =  33°F. 

If  the  fuel  is  just  saturated  at  40°F.,  the  entering  temperature 
of  the  air  and  fuel  would  have  to  be  at  least  40  +  33  =  73°F.  to 
permit  all  the  fuel  to  be  vaporized,  if  no  heat  is  supplied  to  the 
mixture  from  outside. 

The  following  table1  gives  data  of  a  similar  nature  for  various 
fuels.  Column  3  gives  the  temperature  of  the  fuel-air  mixture  at 

1  From  KUTZBACH,  Technical  Note  No.  62.  National  Advisory  Committee  for 
Aeronautics,  1921. 


FUELS  AND  EXPLOSIVE  MIXTURES 


231 


which  the  vapor  of  the  fuel  is  just  saturated;  the  mixture  is 
supposed  to  be  chemically  correct  and  the  pressure  of  the  mixture 
is  atmospheric  pressure.  In  column  4  is  given  the  fall  in  tempera- 
ture of  the  air  and  liquid  fuel  required  to  supply  the  latent  heat 
for  complete  vaporization.1  The  initial  temperature  of  the 
air  must  be  at  least  equal  to  the  sum  of  the  quantities  given 
in  columns  3  and  4;  this  sum  is  given  in  the  last  column. 
SATURATION  TEMPERATURES  OF  AIR-FUEL  MIXTURES 


Fuel 

Boiling 
point, 
deg.  F. 

Saturation 
temperature 
of  the  fuel 
mixture, 
deg.  F. 

Drop  in 
temperature 
due  to 
evaporation, 
deg.  F. 

Minumum 
temperature 
of  the  air  for 
complete 
vaporization, 
deg.  F 

Hexane  

158 

0 

54 

54 

Benzene  . 

176 

23 

54 

77 

Ethyl  alcohol  

172 

72 

198 

270 

Decane  
Naphthalene  

320 

428 

108 
198 

63 

72 

171 
270 

It  is  evident  that  the  temperature  of  the  air-fuel  mixture  with 
decane  or  naphthalene  as  fuel  is  so  high  as  to  reduce  considerably 
the  volumetric  efficiency  and  the  power  of  the  engine  if  all  the 
fuel  enters  in  the  vapor  form. 

Gaseous  Explosions. — In  an  airplane  engine  making  1,800 
revolutions  per  minute,  the  duration  of  the  explosion  should  not 
be  greater  than  the  time  of  one-sixth  of  a  revolution  or  }{  go 
second.  The  possibility  of  employing  a  gasoline  engine  depends 
on  the  possibility  of  carrying  out  the  explosion  process  with  a 
high  degree  of  completeness  in  this  extremely  short  time. 

Explosion  is  a  chemical  reaction  attended  by  the  liberation  of 
a  considerable  amount  of  heat.  It  is  a  combustion  process. 
Combustion  results  from  the  chemical  union  of  a  fuel  with  oxygen 
and  this  union  may  take  place  either  (1)  at  the  place  where  the 
two  are  brought  into  contact  as  with  the  ordinary  gas  burner,  or 
(2)  in  an  intimate  mixture  of  the  two,  as  in  a  bunsen  burner  or 
in  a  gas  engine  cylinder.  Explosive  reaction  can  take  place 
only  with  an  intimate  mixture. 

The  reaction  in  an  intimate  mixture  is  not  necessarily  explosive ; 
for  example,  no  explosion  occurs  in  the  bunsen  burner.  An 

1  Some  of  these  values  are  calculated  by  Kutzbach  from  values  of  latent 
heat  which  are  apparently  too  high. 


232 


THE  AIRPLANE  ENGINE 


explosion  is  always  self -propagating :  that  is,  if  part  of  the  mixture 
is  ignited  the  combustion  will  jgpread  throughout  the  mass  of 
the  mixture.  The  term  "  explosion "  is  commonly  reserved  for 
the  case  where  the  velocity  of  such  propagation  is  high;  but  there 
is  no  definite  line  of  demarcation  between  explosion  and  slow 
burning. 

The  velocity  of  propagation  of  combustion  in  an  explosive 
mixture  depends  on  the  kind  of  fuel,  the  amount  of  oxygen 
present,  the  amount  of  inert  gases  present,  the  temperature, 
pressure,  and  a  number  of  other  factors.  The  strength  of  the 
explosive  mixture  is  the  most  important  factor.  No  explosion 
is  possible  if  the  ratio  of  air  to  fuel  exceeds  certain  limits.  Bunte1 
has  found  the  explosive  limits  for  various  air-fuel  mixtures  at 
atmospheric  pressure  and  temperature  as  given  in  the  following 
table : 

EXPLOSIVE  LIMITS  OP  AIR-FUEL  MIXTURES 


^Fuel 

Ratio  of  air  to  gas  by 
volume 

Theoretical 
ratio  of  air 
to  gas  by 
volume 

Lower  limit, 
air  in 
excess 

Upper  limit, 
gas  in 
excess 

Carbon  monoxide 

5.06 
9.58 
7.06 
28.8 
11.6 
23.4 
24.3 
15.4 
35.7 
36.7 
40.7 

.0.33 
0.50 
0.49 
0.91 
4.23 
5.84 
6.32 
6.81 
12.0 
14.4 
19.4 

2.4 
2.4 
2.4 
11.98 
5.7 
14.4 
14.4 
9.63 
28.41 
36.0 
37.5 

Hydrogen.  .  .         

Water  gas  

Acetylene                            .           .    . 

Coal  gas  

Ethylene  ....                   .... 

Alcohol  

Marsh  gas. 

Ether  

Benzene 

Pentane  

Burrell  and  Gauger2  give  explosive  limits  of  air-gasoline  mixtures 
as  66  and  16  (ratio  of  air  to  gasoline  vapor  by  volume). 

The  above  results  were  obtained  with  mixtures  at  ordinary 
atmospheric  pressures  and  temperatures.  They  show  that  a 
self-propagating  combustion  is  possible  with  most  fuels  where 

1  The  Engineer,  March  28,  1902. 

8  Technical  Paper  150,  U.  S.  Bureau  of  Mines. 


FUELS  AND  EXPLOSIVE  MIXTURES  233 

there  is  a  considerable  excess  present  either  of  air  or  of  fuel. 
These  limits  are  considerably  extended  as  temperature  and  pres- 
sure increase.  For  example,  at  600°C.  it  is  possible  to  explode 
a  mixture  of  CO  with  12  times  its  volume  of  air,  as  compared  with 
5.06  times  at  atmospheric  temperature.  The  presence  of  carbon 
dioxide  in  place  of  some  of  the  excess  air  diminishes  the  explosive 
limits. 

The  temperature  to  which  part,  or  all,  of  the  mixture  must  be 
brought  to  initiate  an  explosion  is  called  the  ignition  temperature. 
This  varies  with  the  fuel,  strength  of  mixture,  the  volume  or  mass 
of  the  mixture  heated,  the  temperature  and  dimensions  of  the 
containing  vessel,  and  the  method  of  ignition.  A  weak  spark, 
although  it  has  a  temperature  much  higher  than  the  ignition 
temperature,  may  fail  to  cause  an  explosion.  It  may  start 
combustion  at  the  place  where  it  passes,  but  the  heat  loss  by 
convection,  conduction  and  radiation  may  be  in  excess  of  the  heat 
of  combustion  and  the  flame  will  fail  to  propagate.  A  sufficient 
duration  of  spark  is  also  necessary.  A  flame  may  ignite  a 
mixture  that  cannot  be  exploded  by  a  spark,  because  it  gives, 
initially,  so  large  a  volume  of  flame  that  the  radiation  loss  to  the 
containing  vessel  does  not  cool  it  below  the  ignition  temperature. 
If  the  whole  mass  is  raised  in  temperature  simultaneously  (as 
by  adiabatic  compression)  the  ignition  temperature  will  be  less 
than  when  part  of  the  mixture  only  is  heated.  This  ignition 
temperature,  with  adiabatic  heating  of  fuel-air  mixtures,  is 
from  about  1,200°F.  for  hydrogen  to  about  1,700°F.  for  carbon 
monoxide.  With  the  usual  gas  engine  fuels  it  falls  between  those 
limits,  the  value  depending  on  the  hydrogen  and  the  neutrals 
present. 

The  ignition  temperature  has  great  importance  as  it  determines 
the  permissible  ratio  of  compression,  and  thereby,  the  limit  of 
efficiency  in  the  engine.  Compression  must  stop  just  short  of 
that  temperature  at  which  ignition  will  occur.  Any  means  for 
increasing  the  cooling  of  the  mixture  during  compression  (such 
as  improved  water  jacketing)  will  permit  a  greater  ratio  of  com- 
pression. Local  heating  of  the  mixture,  as  by  carbon  deposit, 
may  result  in  preignition. 

Combustion  once  started  in  an  explosive  mixture  may  either 
die  out  or  be  propagated.  If  it  once  starts  to  propagate  itself, 
it  is  likely  to  continue  and  there  will  result  an  explosion.  The 
velocity  with  which  the  combustion  is  propagated  increases 


234  THE  AIRPLANE  ENGINE 

progressively  in  all  true  explosions.  In  the  case  of  a  bunsen 
burner  the  velocity  remains  constant  and  the  combustion  is  not 
explosive.  The  flame  in  that  case  is  stationary,  but  as  the  gas  is 
moving  the  flame  is  really  moving  relative  to  the  gas,  in  the 
opposite  direction  and  with  the  same  velocity.  If  the  velocity 
of  the  gas  is  diminished  too  much  by  partly  closing  the  gas 
supply,  the  flame  will  shoot  back,  i.e.,  the  flame  will  travel  more 
rapidly  than  the  gas.  The  flame  remains  at  the  mouth  of  the 
burner  under  considerable  variations  of  gas  velocity  in  the 
burner  because  the  velocity  of  the  mixture  decreases  rapidly  as 
it  issues  from  the  burner,  so  that  there  will  be  some  place,  close  to 
the  burner,  at  which  the  gas  velocity  equals  the  velocity  of  flame 
propagation.  The  flame  will  remain  stationary  at  that  place. 
The  cooling  effect  exerted  by  the  metal  burner  also  reduces  the 
flame  propagation  velocity. 

If  the  velocity  of  the  gas  which  will  just  keep  the  flame  away 
from  the  burner  is  measured,  it  will  give  a  rough  indication 
of  the  velocity  of  flame  propagation  in  the  mixture.  The  results 
will  not  be  very  accurate  because  of  cooling  and  diluting  influ- 
ences of  the  atmosphere.  Experiments  of  that  general  nature 
show  that,  at  atmospheric  temperature  and  pressure,  for  H  and  0, 
the  velocity  of  propagation  is  about  115  ft.  per  second,  and  for 
CO  and  O  about  4J-£  ft.  per  second.  This  is  for  the  combining 
proportions,  which  give  approximately  maximum  velocities. 
With  H  and  air  the  velocity  drops  to  about  10  ft.  per  second 
at  212°F.  For  gasoline-air  mixtures,  at  atmospheric  temper- 
atures, velocities  of  about  3.5  ft.  per  second  and  for  alcohol  about 
3  ft.  per  second  are  realized.  These  results  apply  only  to  linear 
propagation  at  atmospheric  pressure. 

In  a  closed  vessel,  such  as  a  gas  engine  cylinder,  the  conditions 
are  quite  different.  The  propagation,  starting  from  a  point,  is 
spherical;  the  increase  of  temperature  results  in  increase  of 
pressure  and  as  the  flame  spreads  the  unburned  portion  will 
be  compressed  adiabatically  and  will  increase  continually  in 
pressure  and  in  temperature.  As  the  temperature  increases  the 
rate  of  propagation  will  increase.  The  velocity  of  propagation 
will  then  be  continually  accelerated.  The  flame,  moreover,  is 
carried  forward  bodily  by  the  expansion  of  the  burned  portion. 

Experiments  on  explosions  in  closed  vessels  have  determined 
the  time  required  to  reach  maximum  pressure  with  various 
mixtures  exploded  in  vessels  of  various  shapes.  If  the  maximum 


FUELS  AND  EXPLOSIVE  MIXTURES  235 

distance  from  the  ignition  point  to  the  boundary  of  vessel  is 
divided  by  this  time,  the  quotient  gives  a  measure  of  the  average 
rate  of  flame  propagation.  With  illuminating  gas  at  atmospheric 
temperature,  in  a  tube  J-£  m-  m  diameter  and  with  7J^  in.  travel 
of  flame,  this  varies  from  5  to  24  ft.  per  second,  according  to 
the  strength  of  the  mixture.  It  increases  rapidly  with  increased 
initial  temperature;  in  some  cases  as  the  tenth  power  of  the 
absolute  temperature  (=  1,000-fold  for  doubled  temperature). 

With  the  largest  existing  gas  engines  (using  blast-furnace  gas) 
the  available  time  for  a  good  explosion  is  about  ^  sec.  and  the 
maximum  distance  the  flame  must  travel  is  about  IJ^j  ft.;  this 
gives  a  mean  velocity  of  11  ft.  per  second.  The  addition  of  a 
third  igniter  has  sometimes  increased  the  capacity  20  per  cent  and 
shows  that  the  speed  limit  has  been  reached.  Blast-furnace  gas 
consists  mainly  of  CO,  which,  at  low  temperature,  has  a  velocity 
of  propagation  not  greater  than  one-third  that  of  gasoline. 

With  an  airplane  engine  at  1,800  r.p.m.,  the  time  for  explosion 
is  about  Hso  second;  if  the  flame  travels  2  in.  the  mean  velocity 
will  be  33  ft.  per  second.  By  increasing  the  ignition  lead,  still 
more  time  might  be  provided ;  the  speed  of  the  airplane  engine  is 
not  yet  limited  by  the  velocity  of  propagation  of  the  explosion. 
Alcohol  is  slower  so  that  alcohol  engines  could  not  be  run  as  fast 
as  gasoline  engines  if  the  rate  of  propagation  of  the  explosion 
should  ultimately  determine  the  limit  of  speed,  instead  of  valve 
areas  and  inertia  effects  as  at  present. 

The  observed  velocities  of  propagation  in  actual  engines  are 
higher  than  those  which  experiments  with  closed  vessel  indicate. 
This  results  from  another  factor,  turbulence.  The  velocity  with 
which  the  gases  enter  gas-engine  cylinders  is  very  much  higher 
than  the  velocity  of  propagation  of  flame.  With  1-lb.  drop  of 
pressure  into  the  cylinder  and  no  frictional  resistance  the  velocity 
of  the  entering  air  would  be  about  350  ft.  per  second;  with  J4  lb., 
175  ft.;  with  ^{Q  Ib.  about  120  ft.  per  second.  This  gas  velocity 
causes  turbulent  conditions  which  cannot  be  quieted  down  by  the 
time  explosion  starts.  The  propagation  is  not  spherical  but  is 
by  currents  and  eddies  of  burning  gas  which  carry  flame  to  all 
parts  of  the  vessel  more  rapidly  than  is  possible  with  spher- 
ical propagation.  Recent  experimental  work  bears  this  out. 
Dugald  Clerk  found,  in  a  common  gas-engine  cylinder,  that  after 
quieting  down  turbulence,  the  explosion  takes  nearly  three  times 
as  long  as  when  the  usual  conditions  exist.  Experiments  in 


236  THE  AIRPLANE  ENGINE 

closed  vessel  without  stirring  gave  the  time  of  explosion  as  0.13 
sec.;  with  vigorous  stirring  the  time  required  was  only  one-sixth 
as  long. 

Detonation. — During  explosion  in  a  closed  vessel  the  advancing 
flame  sphere  sends  off  compression  waves  which  travel  through 
the  unburned  mixture  with  the  velocity  of  sound  in  that  medium. 
If  the  vessel  is  of  sufficient  dimensions  the  increasing  velocity  of 
the  flame  and  the  continuously  increasing  pressure  and  tempera- 
ture of  the  unburned  mixture  will  result  in  the  formation  of  a 
wave  in  which  the  pressure  will  be  such  as  to  bring  the  mixture 
(adiabatically)  to  the  ignition  temperature.  In  that  case  the 
wave  will  cause  combustion  as  it  moves  on.  The  velocity  of  this 
wave  will  be  greater  than  that  of  sound  because  the  process  is 
not  merely  one  of  wave  transmission  but  of  chemical  reaction 
also.  Investigations  of  the  explosive  wave  show  velocities  of 
the  order  of  magnitude  of  3,000  to  6,000  ft.  per  second  and 
pressures  of  1,000  to  2,000  Ib.  per  square  inch.  These  pressures 
are  destructive  to  engines  and  should  be  avoided. 

The  "detonations"  or  "pinking"  which  are  both  felt  and 
heard  in  engine  cylinders  under  certain  conditions  of  operation 
probably  indicate  either  the  generation  of  an  explosive  wave  or 
breaking  down  of  the  fuel  with  the  liberation  of  free  hydrogen, 
which  explodes  with  extreme  rapidity.  In  such  cases  the  com- 
bustion is  notably  incomplete,  the  exhaust  containing  much 
free  carbon,  and  the  power  and  efficiency  of  the  engine  fall  off. 
Fuels  consisting  of  paraffins  have  a  low-ignition  temperature, 
and  are  readily  detonated.  Fuels  belonging  to  the  aromatic 
group  have  higher  ignition  temperatures  and  can  be  used  with 
higher  compression  pressures  without  detonation. 

The  maximum  pressures  to  which  fuels  can  be  compressed 
without  serious  detonation  have  been  determined  by  Ricardo,1 
who  used  for  that  purpose  a  variable  compression  engine  with 
compact  combustion  space,  central  igniter,  and  other  features 
making  for  maximum  capacity  and  efficiency.  His  results, 
including  the  corresponding  indicated  mean  effective  pressures 
and  thermal  efficiencies,  are  given  in  Table  15.  The  data  for 
toluene,  xylene,  and  acetone  are  for  a  compression  ratio  of  seven, 
which  gives  a  compression  pressure  well  below  their  detonation 
pressures;  it  was  not  considered  desirable  to  go  above  that  com- 
pression ratio  for  hydrocarbon  fuels  on  account  of  the  excessive 

1  The  Automobile  Engineer,  Jan.  and  Feb.,  1921. 


FUELS  AND  EXPLOSIVE  MIXTURES 


237 


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238 


THE  AIRPLANE  ENGINE 


explosion  pressures  reached.  Ricardo  concludes  from  his  investi- 
gations that  detonation  is  less  the  lower  the  rate  of  burning  of  the 
fuel,  and  that  no  fuel  has  a  rate  of  burning  too  low  to  permit  of 
maximum  efficiency  being  obtained  at  the  highest  practicable 
engine  speeds.  All  the  hydrocarbon  fuels  give  the  same  power 
output  and  efficiency,  within  2  per  cent,  for  the  same  ratio  of 
compression  so  long  as  that  compression  is  not  high  enough  to 
produce  detonation.  Figure  169  gives  the  indicated  mean 
effective  pressure  and  indicated  thermal  efficiency  practically 
attainable  in  a  high-grade  engine  with  a  non-detonating  hydro- 
carbon fuel.  With  alcohol  the  high  latent  heat  permits  the  use  of 
a  greater  weight  of  charge  and  consequently  of  increased  output 
as  compared  with  the  hydrocarbons.1 


35      4.0      4.5       5.0       5.5      60      6.5 
Ratio  of  Compression. 

FIG.  169. — Maximum  attainable  m.e.p.  and  thermal  efficiency  in  Otto  cycle 
engine  using  hydrocarbon  fuel. 

The  performance  of  a  mixture  of  hydrocarbons  is  found  to  be 
the  mean  performance  of  the  components.  The  highest  per- 
missible compression  pressure  is  determined  by  the  relative  pro- 
portions of  aromatics,  naphthenes  and  paraffins;  the  smaller  the 
proportion  of  the  last  the  higher  may  be  the  compression  pressure. 
The  heavier  compounds  in  the  paraffin  series  detonate  at  lower 
compression  pressures  than  the  lighter  compounds. 

The  tendency  of  a  fuel  to  detonate  is  expressed  by  Ricardo  in 
terms  of  its  "toluene  value."  The  scale  of  toluene  values  is 
based  on  compression  pressures  at  detonation ;  it  is  0  per  cent  for 
a  selected  standardized  gasoline  detonating  at  a  compression  ratio 
of  4.85,  is  100  per  cent  for  toluene,  and  varies  in  direct  propor- 
tion to  the  change  in  compression  pressure.  Its  value  is  negative 
for  fuels  which  detonate  at  lower  compression  pressures  than 
the  standard  gasoline. 

xThe  vaporization  takes  place  mainly  in  the  cylinder,  during  admission 
and  compression,  and  keeps  down  the  compression  temperature. 


FUELS  AND  EXPLOSIVE  MIXTURES 


239 


The  influence  on  the  detonating  temperature,  and  consequently 
on   the   permissible   compression   pressure,   of  the  addition  of 
toluene  to  a  paraffin  fuel  is  shown  by  the  following  tests.1 
Toluene,  per  cent. ...     0.0       10.0      20.0      30.0      40.0      50.0      60.0 
Compression  ratio...     4.85      5.20      5.57      5.94      6.32      6.67      7.05 
Indicated  m.e.p 132.5     135.4     138.7     142.0     144.9     147.5    150.0 


420 
400 


u    360 

I 

&-360 

o 

i    340 


320 
300 
280 
260 


Brvke- 


Pznna.68°Be  Gaso/im 
Calif.  S9'  Bt  Gasoline 


UJ 

110   i 


100 


1000       1200        1300       1400        1500        1600        1700        1800 

Revo  I  uf  ions  per  Mi n.  •   • 

FIG.  170. — Variation  of  power  of  12-cylinder  Liberty  engine  with  fuel. 

The  effect  of  the  additi9n  of  ethyl  alcohol  is  even  more  marked 
than  that  of  toluene;  only  three-fifths  as  much  alcohol  need  be 
added  to  obtain  the  same  compression  ratios.  The  effect  of  these 
additions  on  engine  capacity  are  indicated  by  the  mean  effective 
pressure  values  of  the  table;  the  increase  in  engine  efficiency  can 
be  seen  from  Fig.  169. 

The  preponderating  importance  of  the  detonating  character- 

1  RICARDO:  Proc.  Royal  Aeronautical  Society,  1920. 


240 


THE  AIRPLANE  ENGINE 


istics  of  a  fuel  has  received  more  recognition  in  England  than  in 
this  country.  Fuels  are  blended  with  toluene,  benzol  or  other 
aromatics  or  naphthenes  so  as  to  give  a  standard  toluene  value. 

This  amounts  to  giving  a 
standard  detonating  com- 
pression pressure  when  the 
fuel  is  used  in  a  standard 
engine.  The  actual  deto- 
nating pressure  and  there- 
fore the  permissible 
compression  ratio  is  largely 
determined  by  the  charac- 
teristics of  the  engine  in 
which  the  fuel  is  used. 
With  a  poorly  shaped  com- 
bustion space,  non-central 
ignition,  and  other  un- 
favorable features,  a  fuel 
will  detonate  at  a  much 
lower  compression  ratio 
than  when  the  conditions 
are  favorable.  In  such  an 
engine,  detonation  may  be 
prevented  by  the  use  of 
overrich  mixtures  and  late 
ignition,  with  a  resulting 
sacrifice  of  both  power  and 
economy. 

Influence  of  Fuel  on 
Capacity.  —  A  comparison 
of  the  power  output  of  a 
Liberty  12  with  two  grades 


FIG 


1000 


1800 


1200          1400         1600 
Revolutions  per  Mm. 

171. — Variation  of  power  of  single-cyl- 
inder Liberty  engine  with  fuel. 


of  gasoline  is  shown  in 
Fig.  170;  the  low-grade 
(59°Be.)  gasoline  falls  off 
rapidly  in  brake  mean  effective  pressure  above  1,500  r.p.m.,  while 
the  68°Be.  gasoline  maintains  its  mean  effective  pressure  well  to 
1,800  r.p.m. 

A  comparison  of  six  fuels  is  shown  in  Fig.  171.  These  fuels 
were  used  in  a  single-cylinder  Liberty  engine.  Their  distillation 
curves  are  given  in  Fig.  172;  these  represent  about  the  full  range  of 


FUELS  AND  EXPLOSIVE  MIXTURES 


241 


commercial  airplane  fuels.  It  will  be  seen  from  Fig.  171  that  the 
total  range  of  power  is  2.8  h.p.  at  1,800  r.p.m.  with  a  maximum 
value  of  37  h.p.  at  that  speed;  this  power  range  is  only  7.6  percent. 


356 


320 


First     10 
Drop 


30      40        50      60       70 
Percentage  Distiffed 
FIG.  172. — Distillation  curves  of  the  fuels  of  Fig.  171. 


90      Dm 
Poinf 


The  tests  of  the  Bureau  of  Mines1  show  the  comparatively 
small  range  in  efficiencies  resulting  from  the  use  of  different 
fuels.  The  fuels  include  straight  refinery,  blended  casing-head, 

CALORIFIC  VALUE,  POWER  DEVELOPED  IN  ENGINE  TESTS,  AND  SPECIFIC 

GRAVITY  OF  VARIOUS  TYPICAL  GASOLINES  FROM  MID-CONTINENT 

AND  EASTERN  FIELDS 


Gravity 

High  calorific 
value  of  gasoline 

Power 
developed, 

T?'    1^1    t                   U"    I* 

v 

sample  was 
obtained 

Process  of 
manufacture 

Specific 

Calories 

B.t.u. 

power 
hours  per 

gravity 

B6. 

per 
gram 

per 
pound 

pound  of 
gasoline 

Mid-Continent  

Cracking  plant 

0.745 

57.9 

11,165 

20,097 

1.345 

Mid-Continent  

"Straight"  refinery 

0.742 

58.7 

11,174 

20,113 

.403 

Mid-Continent  

"Straight"  refinery 

0.733 

61.0 

11,180 

20,124 

.350 

Eastern  

"Straight"  refinery 

0  718 

65  0 

11,187 

20,137 

.405 

Mid-Continent  

"Straight"  refinery 

0.724 

63.4 

11,215 

20,187 

.395 

Mid-Continent  

"Straight"  refinery 

0.727 

62.6 

11,221 

20,198 

.396 

Eastern  

Blended  casing-head 

0  733 

61.0 

11,230 

20,214 

.376 

Eastern  

"Straight"  refinery 

0  724 

63.4 

11,236 

20,225 

.420 

Mid-Continent  

"Straight"  refinery 

0.715 

65.8 

11,250 

20,250 

1.365 

Eastern 

"Straight"  refinery 

0.687 

73.8 

11,315 

20,367 

1.487 

and  cracked  gasoline  with  densities  varying  from  57.9  to  73.80Be\ 
The  accompanying  table  shows  that  the  work  done  per  pound  of 
1  Technical  Paper  163,  U.  S.  Bureau  of  Mines. 


16 


242  THE  AIRPLANE  ENGINE 

gasoline  varied  from  1.345  to  1.487  h.p.h'.,  the  higher  value  being 
generally  obtained  with  the  lighter  gasolines.  The  heats  of 
combustion  also  increased  slightly  as  the  gasoline  became  lighter, 
but  the  total  range  of  higher  heat  values  is  only  slightly  greater 
than  1  per  cent.  Neglecting  this,  and  assuming  alow  heat  value 
of  18,500  B.t.u.  for  all  the  fuels,  the  range  of  efficiencies  is  seen 

,     ,         1.345  X  33,000  X  60 
to  be  from 18  500  X  778     "  =  1^*45  per  cent  to  2^.48  per 

cent.  The  engine  on  which  the  tests  were  made  is  of  low  com- 
pression and  low  efficiency;  the  indications  are  that  the  change  in 
efficiency  with  the  fuel  is  even  less  in  engines  of  higher  efficiency 
such  as  are  used  in  airplane  practice. 

A  mixture  of  alcohol  and  gasoline  has  been  used  in  high- 
compression  aviation  engines.  This  mixture  eliminates  deton- 
ation and  has  good  starting  characteristics.  An  "alcogas" 
tested  at  the  Bureau  of  Standards1  contained  40  per  cent  alcohol, 
35  per  cent  gasoline,  17  per  cent  benzol  and  8  per  cent  of  other 
ingredients.  With  a  compression  ratio  of  5.6,  the  maximum 
power  at  ground  level  was  the  same  as  for  a  high-grade  aviation 
gasoline,  but  as  the  level  increased  the  power  output  became  4  to 
6  per  cent  greater  than  for  gasoline.  The  thermal  efficiency  was 
superior  by  about  15  per  cent;  the  fuel  consumption  was  increased 
on  account  of  the  lower  heat  value  per  pound  of  fuel.  At  a 
compression  ratio  of  7.2  the  power  output  was  increased  by 
about  16  per  cent  as  compared  with  gasoline  at  5.6  compression 
ratio  and  the  thermal  efficiency  increased  about  22  per  cent,  which 
just  offsets  the  lower  heat  of  combustion  of  the  fuel  and  gives 
the  same  weight  of  fuel  per  brake  horse-power  hour  for  both  fuels. 
The  distillation  curve  of  the  alcogas  is  very  flat,  80  per  cent  distill- 
ing off  between  140  and  175°F.;  the  end  point  is  high,  360°F. 

EXPERIMENTAL  DETERMINATION  OF  STRENGTH  OF  MIXTURE 

The  accurate  determination  of  the  ratio  of  air  to  fuel  used  by 
an  engine  requires  the  separate  measurements  of  the  weights  of 
fuel  and  air.  The  weight  of  fuel  is  readily  ascertained,  but 
the  weight  of  air  offers  difficulties.  The  simplest  method  is 
to  connect  the  air  intake  of  the  carburetor,  in  an  airtight  manner, 
with  a  large  box  to  which  air  is  admitted  through  a  standard  or 
calibrated  sharp-edged  orifice.  If  there  is  more  than  one  air 
intake  it  is  better,  if  practicable,  to  enclose  the  whole  carburetor 

1  National  Advisory  Committee  for  Aeronautics,  Report  No.  89. 


FUELS  AND  EXPLOSIVE  MIXTURES 


243 


in  an  airtight  chamber,  connected  with  the  orifice  box.  The  air 
should  enter  the  orifice  quietly,  passing  through  a  honeycomb 
screen  to  eliminate  the  effect  of  wind  or  air  currents. 

The  weight  flowing  through  such  an  orifice  is  given  by  the 
equation  

M  =  l.lF^(p-Po)  XC 

where,  M  is  the  weight  of  air  flowing  per  second  in  pounds. 
F  is  the  area  of  the  orifice  in  square  inches. 
p  is  the  external  atmospheric  pressure,  pounds  per  square 

inch  absolute. 
T   is   the   atmospheric   temperature,    degrees   absolute, 

Fahrenheit. 
Po  is  the  pressure  inside  the  orifice  box,  pounds  per  square 

inch  absolute. 
C  is  a  constant 

The  value  of  C  has  been  determined  by  Durley1  with  great 
accuracy  for  circular  sharp-edged  orifices  in  steel  plates,  0.057 
in.  thick.  The  following  table  gives  his  values  for  the  coeffi- 
cient C: 

COEFFICIENTS  OF  DISCHARGE  FOR  SHARP-EDGED  ORIFICE 


Pressure  difference  on  two  sides  of  orifice,  inches  of  water 


Diameter  of 
orifice,  inches 

1 

2 

3 

4 

5 

KG 

0.603 

0.606 

0.610 

0.613 

0.616 

i^ 

0.602 

0.605 

0.608 

0.610 

0.613 

1.0 

0.607 

0.603 

0.605 

0.606 

0.607 

2.0 

0.600 

0.600 

0.600 

0.600 

0.600 

3.0 

0.599 

0.598 

0.597 

0.596 

0.596 

4.0 

0.598 

0.597 

0.595 

0.594 

0.593 

4.5 

0.598 

0.596 

0.594 

0.593 

0.592 

It  will  be  observed  that  the  coefficient  is  constant  for  a  2-in. 
orifice. 

In  most  cases  it  will  not  be  found  practicable  to  measure  the 
air  in  this  manner.  A  good  approximation  can  be  obtained  from 
a  measurement  of  the  pressure  drop  from  the  mouth  to  the 

i  Trans.  Am.  Soc.  Mech.  Eng.,  1906. 


244 


THE  AIRPLANE  ENGINE 


throat  of  the  choke  or  venturi  tube  of  the  carburetor.  This 
drop  can  be  obtained  by  connecting  a  water  manometer  with  a 
very  small  hole  (^2  in.)  drilled  into  the  smallest  section  of  the 
choke.  Tests  carried  out  on  a  considerable  number  of  carbu- 
retors show  that  the  coefficient  of  discharge,  C,  varies  only 
slightly  with  the  form  and  dimensions  of  the  venturi  tube.  The 
weight  of  air  flowing  through  a  carburetor  of  F  sq.  in.  free  area 
at  the  throat  is  given  by  the  equation 

122-58 


VT 


•Po y-~  _    p.y*  -xc 


i 


where  M ,  p  and  T  have  the  same  meanings  as  for  an  orifice  and 
Po  is  the  pressure  at  the  throat.  The  coefficient  C  varies  from 
0.82  to  0.85  and  may  be  assumed  to  have  the  mean  value  0.84. 


10 

C 

02 

4 

0)       ' 

E    i? 

.t 

,-^- 

—  —  . 

X 

/ 

1 
-2     in 

s 

s* 

<^ 

\ 

/ 

>     IU 

f?  a 

^ 

*s 

,s 

^s, 

v 

^ 

V 

ff. 
0 

/ 

S 

^ 

^ 

/ 

/ 

X 

^ 

^° 

2 

/ 

Xc 

D 

s 

S    4 

Ju      7 

x 

x 

s 

<£:    c 

n 

s 

X 

20         18         16         14 

Ratio  of  Air  to  Gasoline 

FIG.  173. — Composition  of  the  exhaust  gases  from  a  gasoline  engine. 

Still  another  method  is  available  if  actual  air  and  fuel  measure- 
ments are  impracticable.  The  investigations  of  Watson,  on 
automobile  engines,  have  shown  that  the  composition  of  the 
exhaust  gases  varies  in  a  regular  manner  with  the  strength  of 
the  mixture  of  air  and  gasoline  admitted  to  the  cylinder.  His 
results  are  shown  graphically  in  Fig.  173.  With  a  chemically 
perfect  mixture  of  about  14.5  parts  of  air  to  one  of  gasoline  the 
exhaust  gases  contain  about  13  per  cent  of  C02  by  volume  and 
about  0.5  per  cent  each  of  02  and  CO.  If  the  air  is  present  in 
excess  (weaker  mixture)  there  is  more  free  02  and  less  C02  in  the 
exhaust;  if  the  mixture  is  richer,  the  free  02  disappears  and  the 
amount  of  CO  increases  while  the  C02  decreases.  All  that  is 
necessary  for  the  test  is  an  Orsat  or  other  volumetric  gas-analysis 
apparatus  and  the  determination  of  the  C02  and  02  content,  or 
in  case  no  02  is  present,  the  CO2  and  CO  content. 


CHAPTER  X 


THE  CARBURETOR 

An  ideal  explosive  mixture  arriving  at  the  intake  manifold 
of  an  engine  should  have  the  following  characteristics:  (1) 
it  should  be  homogeneous  throughout,  (2)  it  should  be  of  the 
composition  or  strength  to  develop  maximum  economy  under 
each  condition  of  engine  operation,  and  (3)  it  should  permit  of  the 
development  of  the  maximum  possible  power. 

In  a  stationary  constant-speed  engine,  in  which  engine  torque 
alone  is  variable,  these  results  might  be  approximated  by  the 
use  of  an  injection  valve,  under  the  control  of  the  governor, 
spraying  finely  atomized  fuel  into  the  current  of  air  going  to 
the  cylinders.  In  an  automobile  engine,  with  both  engine  torque 
and  speed  variable,  this  simple  injection  method  cannot  give 
satisfactory  results.  In  the  airplane  engine  with  the  three 
main  variables  of  torque, 

speed  and  air  density,  the  |    ToEnglne 

problem  is  even  more  com- 
plicated. For  such  engines 
the  explosive  mixture  is 
formed  by  the  use  of  a 
carburetor. 

A  carburetor  is  a  device 
in  which  part  or  all  of  the 
air  going  to  the  engine 
passes  through  a  restricted 
passage,  thereby  acquiring 
velocity  with  consequent 
fall  of  pressure;  the  fuel  is 
sucked  into  the  current  of 
air  in  an  amount  which 
varies  with  the  pressure 

drop.  In  the  simplified  standard  form  of  carburetor  shown  in 
Fig.  174,  air  flows  through  the  restricted  " choke,"  C,  and  creates 
a  partial  vacuum.  Gasoline  is  maintained  at  a  constant  level  in 
the  float  chamber  by  the  action  of  the  float,  F,  which  controls  the 
position  of  the  needle  valve,  V,  past  which  the  gasoline  enters. 
As  the  float  chamber  is  open  to  the  atmosphere  the  level  of 

245 


Air 


FIG.  174. — Diagram  of  simple  carburetor. 


246  THE  AIRPLANE  ENGINE 

gasoline  in  the  nozzle  or  jet,  J,  will  be  the  same  as  that  in  the 
float  chamber  so  long  as  the  engine  is  not  operating.  The  dis- 
charge orifice  of  the  nozzle  is  placed  higher  than  the  gasoline 
level  in  the  float  chamber  to  prevent  overflow  of  the  gasoline 
into  the  air  passage  when  the  engine  is  standing  in  such  position 
as  to  incline  the  carburetor  at  a  moderate  angle  to  the  position 
shown  in  the  figure.  When  air  is  drawn  through  the  carburetor, 
increasing  reduction  of  pressure  at  C,  resulting  from  increasing 
velocity  of  the  air,  will  give  an  increasing  head  on  the  gasoline 
and  will  cause  an  increasing  weight  flow  of  the  fuel.  The 
mixture  of  air  and  fuel  will  be  of  constant  strength  if  the  weight 
of  gasoline  discharged  by  the  jet  is  directly  proportional  to  the 
weight  of  air  flowing  through  the  choke.  The  actual  strength 
of  the  mixture  whether  constant  or  not  is  controlled  by  the  size 
of  the  gasoline  jet. 

A  carburetor  built  as  in  Fig.  174  would  not  discharge  a  mix- 
ture of  constant  strength  for  all  rates  of  air  flow,  nor  is  such 
constancy  desirable.  It  is  common  experience  that  the  mixture 
delivered  to  the  engine  should  be  richer  at  very  light  loads 
(idling)  than  for  heavier  loads,  and  also  that  it  should  be  richer 
for  maximum  power  than  for  maximum  economy.  A  satis- 
factory carburetor  should  vary  the  strength  of  the  mixture  so  as 
to  maintain  the  desired  strength  under  all  conditions  of  operation 
of  the  engine. 

A  study  of  the  action  of  a  carburetor  requires  a  knowledge  of 
the  laws  of  flow  of  gases  and  liquids  through  such  passages  as 
are  found  in  carburetors.  The  more  important  results  of 
experiment  on  such  flow  are  given  in  the  following  pages. 

Theoretical  Flow  of  Air  through  a  Constricted  Tube. — When 
air  is  flowing  steadily  through  a  tube  whose  cross-section  varies, 
the  weight  and  the  total  energy  passing  each  section  of  the  tube 
per  second  are  constant. 

Let  W  =  Weight  of  air  passing  in  pounds  per  second. 

p  =  The  air  pressure  in  pounds  per  square  foot  absolute. 
P  —  The  air  pressure  in  pounds  per  square  inch  absolute. 
v  =  The  specific  volume  of  the  air  in  cubic  feet  per  pound. 
V  =  The  velocity  of  the  air  in  feet  per  second. 
T  =  The  absolute  temperature  of  the  air  in  degrees  Fahrenheit. 
7  =  The  internal  energy  of  the  air  per  pound  in  foot-pounds. 
A  =  The  cross-section  of  the  tube  in  square  feet. 
a  =  The  cross-section  of  the  tube  in  square  inches. 
q  —  Gravitational  acceleration  =  32.16  feet  per  second  per  second. 


THE  CARBURETOR  247 

The  total  energy  passing  any  cross-section  with  unit  mass  of  air 
is  the  sum  of  the  internal  energy  I,  the  displacement  work  pv, 
and  the  kinetic  energy  V2/2g  at  that  section.  Assuming  no 
heat  transfer  through  the  tube  the  total  energy  at  1  (Fig.  175) 
can  be  written  equal  to  that  at  2. 

/i  +  Pi  v,  +         =  72  +  p2v2  +      *2  (1) 


FIG.  175. — Venturi  tube  of  optimum  proportions. 

Air  is  practically  a  perfect  gas.     If  the  expansion  is  without 
eddies  or  friction  and  without  transfer  of  heat  (adiabatic) 

/0v 


n-   1 

_  specific  heat  at  constant  pressure  _ 
~  specific  heat  at  constant  volume 
Furthermore,  with  adiabatic  expansion 

Substituting  from  equations  (2)  and  (3)  in  equation  (1)  there 
may  be  obtained  the  equation 

TV_Zi2=^JL          r          (f*\^*\  (4) 

2g   "  2g       n-l  plVl  I          \pj     J 

The  weight  flow  past  any  section  is  equal  to  the  volume  passing 
that  section  per  second  divided  by  the  specific  volume,  or, 


V 

Since  the  weight  flow  is  constant  at  all  sections 


and  substituting  from  equation  (3) 


248  THE  AIRPLANE  ENGINE 

Substituting  this  value  of  FI  in  equations  (4)  and  (5) 


F*  = 


and 


Pi/  - 


(7) 


W  =  A,   ^ 


Pl 


n-1 


1  -  p 

XPi 


•We5 


(8) 


If  the  section  ^.i  is  taken  just  outside  the  tube  where  the  cross- 
section  may  be  regarded  as  infinite  and  the  air  velocity  zero, 
these  last  equations  become 


(9) 


and 


(10) 


In  the  use  of  equation  (10)  the  specific  volume,  v\t  has  to  be  deter- 
mined from  the  known  pressure,  p\,  and  absolute  temperature,  TI, 
by  the  gas  equation,  p&i  =  RTi  =  53.347Y  Furthermore, 
it  is  practically  most  convenient  to  deal  with  pressures  in  pounds 
per  square  inch,  P,  and  with  areas  in  square  inches,  a.  Substitut- 
ing the  numerical  values  of  g  and  n,  substituting  P  and  a  for  p 
and  A,  and  substituting  53.3477i/pi  for  v\,  equation  (10)  becomes 


W 


//PA1-422       /P2\i. 

Vw      •  IK) 


This  equation  is  difficult  to  use  when  the  desire  dg  weight  flow,  W,  is 
known  and  the  pressure  drop  is  required.  The  curves  of  Fig.  176 
are  based  on  this  equation.  The  ordinates  are  weight  flows  per 
square  inch  of  area  per  minute,  and  the  abscissae  are  pressure 
drops  measured  in  inches  of  water.  One  pound  per  square  inch 
equals  27.70  inches  of  water.  An  initial  temperature  of  60°F. 


THE  CARBURETOR 


249 


(520°  absolute)   is  assumed.     The  separate  curves  are  for  the 
initial  pressures,  PI,  marked  on  them. 

Equations  (8)  and  (10)  have  a  limit  to  their  range  of  applica- 
tion. If  air,  initially  of  pressure  pi  and  specific  volume  vit  flows 
through  a  tube  the  smallest  section  of  which  is  A2,  the  weight 
flow  varies  with  the  pressure,  pz,  at  that  section.  The  weight 


20  40  60  80 

Pressure   Drop  in   Inches  of  Water 

FIG.  176. — Weight  flow  of  air. 


100 


120 


flow  will  be  found  from  the  equation  to  reach  a  maximum  value  as 
Pz  diminishes  to  a  certain  critical  value,  and  then  will  apparently 
diminish  as  pz  is  still  further  reduced.  This  critical  pressure 
occurs  when 


-  - 

pi 


n 


250  THE  AIRPLANE  ENGINE 

That  is,  the  maximum  weight  flow  will  occur  when  the  pressure  at 
the  smallest  cross-section  of  the  tube  is  53  per  cent  of  the  initial 
or  maximum  pressure.  It  is  found  by  experiment  that  the 
pressure  at  the  smallest  cross-section  is  never  less  than  this 
amount,  and  that  it  remains  at  that  exact  value  so  long  as  the 
pressure  on  the  downstream  side  is  equal  to  or  less  than  that 
pressure.  The  weight  flow  through  a  frictionless  tube  is  deter- 
mined by  the  area  of  the  smallest  cross-section  and  cannot  be 
increased  by  decreasing  the  pressure  on  the  downstream  side  of 
that  section  below  the  critical  pressure.  In  equations  (8)  and 
(10)  p%  can  never  have  a  value  lower  than  0.53pi. 

For  very  small  pressure  ranges,  for  example  when  (pi  —  p2)  is 
equal  to  or  less  than  1  per  cent  of  pif  the  expansion  of  the  air 
resulting  from  the  pressure  drop  is  so  small  as  to  be  negligible 
and  the  flow  may  be  assumed  to  follow  the  simpler  laws  of  flow 
of  incompressible  fluids.  In  this  case  I\  =  72,  and  vi  =  v2)  and 
equation  (1)  becomes 

V22       V,2 


If  the  initial  velocity  is  zero 
y  2 


(12) 


or  V  is  proportional  to  \/p1  —  p2,  and  since  the  weight  flow  is 
proportional  to  V  (with  constant  cross-section  and  constant  air 
density), 


W  =  Kpi  -  p*  (13) 

With  air  at  atmospheric  pressure,  the  error  resulting  from  the 
use  of  this  equation  would  be  about  —2.3  per  cent  for  1  Ib.  per 
square  inch  pressure  drop,  and  is  roughly  proportional  to  the 
pressure  drop  for  small  pressure  drops.  Equation  (12)  shows 
that  the  velocity  and  therefore  the  weight  of  air  flowing  is  pro- 
portional to  the  square  root  of  the  pressure  drop  so  long  as  the  pres- 
sure drop  is  small. 

The  actual  flow  of  air  through  a  constricted  tube  is  found  to 
be  less  than  the  amount  indicated  by  equation  (10).  Actual 
flow  is  always  accompanied  by  frictional  resistance  and  the 
formation  of  eddies.  The  ratio  of  the  actual  flow  to  the  theo- 
retical flow  of  equation  (10)  is  called  the  coefficient  of  discharge 
of  the  tube  and  its  value  has  to  be  determined  by  experiment. 

The  choke  of  a  carburetor  is  usually  of  the  general  form  shown 


THE  CARBURETOR  251 

in  Fig.  175,  that  is,  it  consists  of  three  parts:  (1)  a  converging 
entrance;  (2)  a  throat;  and  (3)  a  diverging  discharge;  such  a 
tube  is  generally  described  as  a  venturi  tube.  With  equal  areas 
at  the  entrance  and  exit,  the  pressure  drop  from  the  entrance  to 
the  throat  would  be  entirely  regained  at  the  exit,  if  the  air  flow 
were  frictionless  and  eddyless.  In  actual  carburetors,  the 
pressures  at  discharge  will  be  less  than  that  at  entrance,  and  the 
difference  will  depend  on  the  velocity  of  the  air,  the  "stream- 
lining" of  the  passage,  the  degree  of  obstruction  offered  by  the 
gasoline  jet,  and  the  weight  of  gasoline  carried  by  the  air.  The 
total  pressure  drop  in  the  venturi  is  important  in  determining 
the  volumetric  efficiency  and  capacity  of  the  engine;  to  develop 
maximum  power  the  charge  should  enter  the  cylinder  with  the 
maximum  possible  density.  Loss  of  pressure  in  the  carburetor 
is  a  direct  source  of  loss  of  power  in  the  engine. 

The  results  of  published  tests  on  comparatively  large  venturi 
tubes,  with  straight  axes,  and  without  obstruction  at  throat  or 
entrance,  show  discharge  coefficients  /varying  from  0.94  to  0.99 
for  cases  where  AI  =  AS)  and  A2  is  equal  to  or  less  than  0.5  A\. 
In  these  tubes  it  is  found  that,  for  minimum  friction  and  eddy  loss, 
the  included  angle  for  the  converging  entrance  should  not  exceed 
30  deg.,  and  the  diverging  discharge  tube  should  have  an  included 
angle  between  5  deg.  and  7.5  deg.;  these  should  be  joined  to  a 
short  cylindrical  throat  by  well  rounded  junctions.  Figure  175 
shows  a  venturi  tube  of  these  optimum  proportions. 

Such  optimum  proportions  are  generally  not  practicable  for 
airplanes.  Considerations  of  space  available  make  it  necessary 
to  modify  the  entrance  by  curving  its  axis,  and  force  the  adoption 
of  larger  included  angles.  Furthermore,  the  air  passage  is 
obstructed  by  the  gasoline  jet  and  its  supporting  bosses,  and,  in 
many  cases,  by  the  throttle  valve.  All  these  factors  will  cause 
a  diminution  in  the  discharge  coefficient  and  an  increase  in  the 
pressure  loss.  An  investigation  at  the  Bureau  of  -Standards1 
gives  data  on  certain  carburetors  which  were  designed  for  the 
Liberty  engine.  The  air  passages  of  these  carburetors  are 
shown  in  Fig.  177.  The  tests  were  made  with  various  air 
densities  (corresponding  to  different  altitudes),  and  both  with  and 
without  fuel  admission.  Figure  178  shows  the  coefficient  of 
discharge;  Fig.  179  the  ratio  of  the  exit  to  the  entrance  pressure, 

1  P.  S.  TICE:  National  Advisory  Committee  for  Aeronautics,  4th  annual 
report,  pp.  608-615. 


252 


THE  AIRPLANE  ENGINE 


for  both  carburetors,  with  air  of  750  mm.  pressure  and  with 
various  weights  of  air  flowing.  Figure  180  shows  the  pressure 
recovery  ratio  for  the  Zenith  carburetor,  with  various  air  densities, 
and  both  with  and  without  fuel  admission  to  the  air.  The 
conclusions  derived  from  these  tests  are  as  follows: 


B 
FIG.  177. — Zenith  (A)  and  [Stewart- Warner  (fi)  carburetors. 

1.  The  coefficient  of  discharge  for  the  carburetor  passages 
tested  has  an  almost  constant  and  maximum  value  for  effective 
throat  velocities  greater  than  about  150  ft.  per  second. 


20  30  40 

P,-P2( Inches  of  Water) 


50 


FIG.  178. — Venturi  discharge  coefficients  for  Zenith  (A)   and~Stewart-Warner 

(.B)  carburetors. 

2.  The  value  of  the  coefficient  of  discharge  for  the  carburetor 
passages  tested  lies  between  0.82  and  0.85,  under  service  con- 
ditions. These  values  are  probably  typical  of  reasonably  well 
formed  passages  of  similar  type. 


THE  CARBURETOR 


253 


3.  The  coefficient  of  discharge  for  carburetor  passages  of  this 
type  is  apparently  only  slightly  modified  as  a  result  of  consider- 


^ 

^ 

•>< 

.ft"* 

X 

\ 

"S 

%, 

tC 

^ 

*J 

^ 

•-P 
o 
a: 
Foe 

^ 

<s> 
\ 

\ 

\ 

\ 

E 

\ 

1 

\ 

0.96Q 

02 

0 

04 

a 

)& 

0. 

i 

0. 

16 

0. 

20 

Air-Lb.  per  Sec.  per  Sq.  In.  of  Throat  Area 

FIG.  179. — Pressure  drops  at  partial  loads  in  Zenith  (A)  and  Stewart- Warner 

(B)  carburetors. 

able  changes  in  passage  form,  with  respect  to  angles  of  entrance 
and  exit. 


1.00 


0.99 


a:  0.98 


£0.91 


0.96 


"X 

5 

^v 

A  =  Air  only  at    75 
B=Air    »      »     55 
C-Air     «      »     2,1 
D=  Air  and  fuel  at  15 
csAsr  **     "    *   5i 
f=A/-r  „    „    „  31 

Omm. 
0 

•o 

0 
0 
-Q 

•    M 

\N 

S^ 

^ 

\ 

\ 

L                                 X 

NX 

> 

^ 

\ 

\ 

S3    ' 

^< 

V 

\\ 

£\ 

\\ 

\\ 

V 

\\ 

V 

\ 
\ 

\ 

s 

\ 

\ 

\  \ 
\ 

^ 

\ 

s 

1 

\ 

\ 

\ 

\ 

0.04  0.08  0.12  0.16  0.20 

Air  -  Lb.  per  Sec.  per  Sq.  In.  of  Throat  Area 

FIG.  180. — Pressure  drop  through  a  Zenith  carburetor  as  affected  by  air  density 
and  the  injection  of  fuel. 


4.  The  coefficient  of  discharge  for  a  carburetor  passage  is 
practically  unaffected  by  wide  variations  in  atmospheric  density 


254  THE  AIRPLANE  ENGINE 

(less  than  1  per  cent  maximum  variation  between  the  density 
limits  of  0.075  and  0.035  Ib.  per  cubic  foot). 

5.  The  coefficient  of  discharge  for  a  carburetor  passage  is 
practically  unaffected  by  the  introduction  of  fuel  to  the  air 
stream   (fuel  discharge  introduces  irregularities  not  to  exceed 
plus  or  minus  1  per  cent). 

6.  The  pressure  loss  in  the  carburetor  outlet  changes  with 
the  turbulence  or  internal  motion  of  the  air  stream. 

7.  The  pressure  loss  in  the  carburetor  outlet  changes  with  the 
quantity  of  fuel  admitted  to  the  air  stream,  and  with  the  method 
of  dividing  the  fuel  by  spraying. 

Pulsating  Flow.  —  The  previous  discussion  relates  to  steady 
flow  of  air  through  the  choke.  In  actual  operation  the  flow  is 
pulsating;  each  carburetor  usually  supplies  three  or  four  cylinders. 
With  a  maximum  of  four  cylinders  the  carburetor  will  be  supply- 
ing one  cylinder  only  at  any  instant.  The  flow  of  the  air  through 
the  carburetor  is  determined  by  the  velocity  of  the  piston  in  the 
cylinder  to  which  the  air  is  going.  As  this  velocity  is  zero  at  the 
ends  of  the  stroke  and  a  maximum  at  midstroke,  the  variation  in 
velocity  of  flow  through  the  carburetor  would  be  considerable 
were  it  not  for  the  steadying  effect  of  the  intake  manifold. 
The  volume  interposed  between  the  carburetor  and  cylinder 
acts  as  an  equalizing  device  and  cuts  down  the  pressure  pulsations 
at  the  exit  of  the  carburetor.  Tests  made  in  England  and 
at  the  Bureau  of  Standards1  show  that  for  a  given  weight  of 
air  flowing  under  pulsating  discharge  the  coefficient  of  discharge 
of  the  carburetor  (as  determined  from  pressure  measurements 
at  the  throat),  the  pressure  recovery  ratio,  and  the  strength  of 
the  mixture  are  practically  the  same  as  for  steady  flow. 

The  Flow  of  Fuel  through  a  Nozzle  or  Jet.  —  The  flow  of  a 
liquid  through  an  orifice  is  given  by  the  expression  V  =  C\/2gh  , 
where  V  is  the  velocity  of  flow,  C  a  coefficient,  and  h  the  head 
under  which  the  flow  is  occurring.  This  expression  becomes 


where  W  =  Weight  of  liquid  discharged  in  pounds  per  minute. 
a   =  Area  of  passage  in  square  inches. 
s    =  Specific  gravity  of  the  liquid  (referred  to  water  at 

60°F.) 
h    =  Head  or  pressure  drop  across  the  jet  expressed  in 

inches  of  water. 
1  National  Advisory  Committee  for  Aeronautics,  4th  annual  report,  p.  616. 


THE  CARBURETOR 


255 


The  coefficient  C  includes  losses  due  to  skin  friction,  fluid  friction, 
contraction,  and  end  effects.  Its  value  varies  with  the  head,  h, 
with  change  in  shape  of  the  entrance  to  the  jet,  with  change  in 
ratio  of  length,  L,  to  diameter,  D,  of  the  passage,  and  with  the 
viscosity  of  the  fuel. 

Investigations  by  Tice1  on  the  flow  through  jets  show  the  in- 
fluence of  these  different  factors  on  the  value  of  C.  The  effect 
of  the  alteration  of  the  shape  of  the  jet  entrance  from  square  to 
chamfered  is  shown  in  Fig.  181.  The  diameter  and  length 
are  the  same  for  both  jets.  The  major  effect  of  the  chamfering 
is  to  reduce  the  contraction  of  the  stream  in  the  entrance,  in 
this  case,  at  heads  above  2  in.  in  water.  While  the  coefficient, 


gxu 

c 


.0.4 


O.Z 


a__a 

Squar 


re  Chamfered 


EFFECT  OF  CHAMFERING  ENDS 

OF  PASSAGE 

Includedanqle  of  chamfer  =      60° 
Depth             f       "        =  0.008"' 
Diameter  of  passage      =  0.040 
Length      »        »            -  0.4016' 
Lenqth  •+  depth              =10.04 
I    ^    I L_J 1 1 L_ 


8          12         16         eo 

Head  =h (Inches  of  Water) 


FIG.  181. — Discharge  coefficients  of  square  and  chamfered  jets. 

C,  has  considerably  higher  values  with  increase  of  h  with  the 
entrance  chamfered  in  this  way,  it  will  be  noted  also  that  its 
value  varies  through  wider  limits.  Chamfering  has  the  very 
practical  advantage  in  carburetor  manufacture,  that  the  angle 
and  depth  of  the  chamfer,  within  comparatively  wide  limits, 
have  an  almost  negligible  effect  on  the  discharge;  while,  on  the 
other  hand,  small  departures  from  truth  in  the  making  of  sharp 
square  edges  result  in  wide  variations  in  the  discharge.  This, 
together  with  the  great  difficulty  of  producing  duplicate  parts  hav- 
ing square  edges  free  from  burr,  practically  rules  out  the  square 
edge  for  carburetor  metering  passages. 

Within  the  range  of  metering  passage  diameters  used  in  general 
carburetor  practice,  it  is  found  that  the  value  of  C  increases  with 
increase  of  D  (Fig.  182). 

The  effect  upon  C  of  change  in  the  ratio  L:D  is  brought  out 
in  Figs.  183  and  184.  In  the  former,  C  is  plotted  against  h  for 

lLoc.  cit.,  p.  603 


256 


THE  AIRPLANE  ENGINE 


several  values  of  L:D  with  D  a  constant.     In  Fig.  184,  C  is 
plotted    against    L:D,    each    curve    being  representative   of  a 
constant  value  for  h. 
i.o 


o.8 


0.6 


0.4 


0.2 


EFFECT  ON-C  OF  CHANGE  IM-0 
-^-D  SUBSTANTIALLY  CONSTANT 


Submerqed  Orifice-Chamfered  Ends 
D'*         L"         L^D    T°C      I 


A  =0.0327  0.4041  12.36    23.25 

B*  0.0350       »  11.54    2380 

C  =  0.0373       "  10.84   24.10 

D  =  0.0395       »  10. 23    24.70 

£=  0.0310  0.0150  0.484  27.55 

f  =  0.0357        »  0.42024.45 


12  16  20  24 

Head  =  h(  Inches  of  Water) 


28 


FIG.  182. — Influence  of  diameter  on  the  discharge  coefficients  of  jets. 

1.0 


o0.8 

JC 


0.2 


•  EFFECT  ON-C  OF  CHANGE  IN-L+D 

WITH  D  CONSTANT 
•Submerged  Orifice -Chamfered  Ends 
D"       L"     L*D     T°C 

-  A  =  0.0351   0.406    11.31     21.10 
B=      "      0.200    5.60     26.60 

.  C  =      "      0.100    2.80    23.80 
D=       »      0.015    0.42    24.40 
l        I        I        I        I        I 


1Z  16  20  24 

Head  =  h  (inches  of  Water) 


28 


FIG.  183. — Influence  of  the  ratio  of  length  to  diameter  on  the  discharge  coeffici- 
ents of  jets. 


VALUE  OF  "L-rD  AT  VARIOUS  HEADS 
Submerqeol  Rassaqe  -Chamfered 
0  =  0.0357"  | 


10 
Diameter 


14 


4  6 

L/D  =  Length 

FIG.  184. — Influence  of  the  liquid  head  on  the  discharge  coefficients  of  jets. 

A  change  in  temperature,  T,  affects  the  discharge  from  a  passage 
in  two  ways — through  its  influence  on  the  density,  s,  and  through 


THE  CARBURETOR 


257 


the  change  in  fluidity.  For  ordinary  variations  in  T,  the  change 
in  s  is  comparatively  small  and  has  very  slight  influence  on  the 
discharge.  The  curves  A,  B  and  C,  in  Fig.  185,  for  gasoline 
discharged  from  a  jet  at  three  temperatures,  expresses  the  order 


0.8 


0.6 


0.4 


r  —  ' 

c% 

,B 

& 

*.  — 

&=£ 

->H 

p^ 

o 

,2 

^^, 

^ 

EF 

FECT  ON  -C  OF  CHANGE  IN-T 
Gasoline  Flowinq 
jmerqed  Orifice  -ChamYered  End 
ID"      L"     L-D    T°C 
=  00344  0.407  11.83    9.85  0. 

±    »       »      >.    mo  o. 

=      »          *              29.80  0. 
Free  Orifice  -Square  Ends 
=  0.042    0.005    0.119  24.50 
=  0.020    0.0,05    0.250  2+50  & 

& 

f 

bu 
A 

t 

/ 

B 
C 

W 

tf9 

D 

£ 

4.00 

12  14  -16 

Head  =  h  (Inches  of  Water) 


ZO 


24 


FIG.  185. — Influence  of  temperature  on  the  discharge  coefficients  of  jets. 

of  magnitude  of  the  effect  upon  C  of  change  in  fluidity  resulting 
from  change  in  T.  These  results  are  for  a  comparatively  long 
passage,  in  which  this  effect  is  much  greater  than  with  the  smaller 
values  for  L:D  found  in  carburetor  practice.  The  curves  D  and 


500 

i 

0400 

c 

-4- 

§300 
!!r200 

-4- 
'i5 

C 
100 

0 

pd 

^ 

-V 

. 

X 

X 

2 

r 

x 

^J^ 

^ 

oiS> 

X 

X 

GAS 
NO. 

SP. 
6R. 

X 

^ 

V 

^ 

x 

*y 

X 

/  =0.t80 
2=0.694 
3=0699^ 

X 

s 

^1 

p 

J 

X 

$ 

X 

4=0.702 
5=0.726 

^ 

^r 

s 

^ 

<^ 

? 

^x 

s 

6=0.722 
7=0.7/7 

& 

J^O^- 

^" 

^x^ 

.X 

X 

9=0.748 
10--0.813- 

g^ 

fo- 

> 

^x 

R 

ELATIONSHIP  BETWEEN  FLUID1P 
AND  TEMPERATURE 
FOR 
everal  samples  of  special  aviati 
isoline  and  one  commercial  grade 

f 

"p 

^ 

^ 

^ 

cj-^rl 

gc 

J(9) 

'"" 

.—  o-" 

J 

•)               20               40              60               80              100             120 

140 

Temperature,  Deg.  C. 
FIG.  186. — Variation  of  the  fluidity  of  liquid  fuels  with  temperature. 


E  are  for  sharp-edged  orifices;  they  show  great  constancy  of  C 
with  variation  both  of  h  and  of  T7:  a  change  in  T  from  24.5°C. 
to  4°C.  shows  no  appreciable  change  in  C  at  any  value  of  h. 


17 


258 


THE  AIRPLANE  ENGINE 


The  fluidity  of  a  liquid  is  the  reciprocal  of  its  viscosity.  The 
variation  of  the  fluidity  of  aviation  engine  fuels  with  temperature 
has  been  investigated  by  Herschel,1  who  finds  the  results  shown 
in  Fig.  186.  The  value  of  C  for  a  jet  will  increase  as  the  tem- 
perature and,  therefore,  the  fluidity  of  the  fuel  increases.  There 
is  no  fixed  relation  between  the  densities  and  fluidities  of  different 
fuels;  a  change  of  fuel  will  ordinarily  result  in  a  change  in  C. 

Mixture  .Characteristics  of  a  Carburetor  with  Constant  Air 
Density. — It  has  been  shown  by  equation  (13)  that,  for  moderate 
pressure  drops  in  the  choke,  the  theoretical  air  flow,  W,  is  sensibly 
proportional  to  the  square  root  of  the  pressure  drop,  and  with  a 
constant  coefficient  of  discharge  this  means  that  the  actual 
air  flow  follows  the  same  law.  It  has  been  further  shown  that 
with  discharge  through  a  sharp-edged  orifice  the  flow  of  the 
fuel  follows  the  same  law.  Consequently,  it  would  seem  possible 
to  construct  a  carburetor  in  which  the  air-fuel  ratio  would  remain 
constant  for  moderate  air  flows.  Actual  carburetor  constructions 
do  not,  however,  employ  sharp-edged  orifices  on  account  of  the 
production  difficulties  already  mentioned.  Furthermore,  the  air 
flow  does  not  increase  as  rapidly  as  the  square  root  of  the  pressure 

drop — for  example,  in  Fig. 
176,  with  air  initially  at 
14.7  Ib.  per  sq.  in.  pressure, 
as  the  pressure  drop  in- 
creases from  10  in.  to  40 
in.  of  water,  the  weight 
flow  instead  of  doubling 
increases  only  from  6.56  to 
12.6  or  1.92  times.  At  the 
same  time,  using  the  stand- 
ard form  of  chamfered  jet, 
as  shown  in  Fig.  181,  the 
coefficient  of  discharge  in- 
creases and  thereby  in- 
creases the  flow  of  fluid  more  than  two  fold.  The  mixture  will 
therefore  increase  in  richness  as  the  load  increases.  To  offset 
this  increase,  the  structure  of  Fig.  174  is  modified  in  all  com- 
mercial carburetors.  These  modifications  are  extremely  diverse 
in  character  and  can  be  such  as  to  produce  a  constant  mixture 


10 


FIG.  187. — Variation  of  mixture  strength 
with  load  in  Zenith,  Stewart- Warner  and 
Stromberg  carbureters. 


Bureau  of  Standards  Technologic  Paper  No.  125. 


THE  CARBURETOR  259 

or  almost  any  desired  variation  of  mixture  with  load.  Some  of 
these  constructions  will  be  considered  later. 

The  results  of  tests  on  three  special  airplane  carburetors  at 
standard  air  density,  shown  in  Fig.  187,  are  characteristic  of  the 
methods  of  variation  of  the  air-fuel  ratio  with  the  load  in  actual 
carburetors.  The  Zenith  carburetor  shows  a  very  constant  mix- 
ture; the  other  two  show  enrichment  of  the  mixture  with  dimin- 
ishing load,  a  characteristic  exactly  opposite  to  that  of  the  simple 
carburetor  of  Fig.  174.  The  actual  value  of  the  air-fuel  ratio 
depends  on  the  size  of  the  fuel  orifice  and  is  not  characteristic 
of  the  type  of  construction. 

Mixture  Characteristics  of  Carburetor  with  Variable  Air 
Density.  —  When  the  air  density  changes,  as  a  result  of  change 
of  air  pressure  and  temperature  during  the  ascent  of  an  airplane, 
a  new  disturbing  element  is  introduced  into  the  behavior  of  the 
carburetor.  With  level  flight  and  wide-open  throttle,  the  engine 
speed  may  be  assumed,  as  a  first  approximation,  to  be  constant  at 
all  altitudes;  this  is  not  the  case  since  the  engine  speed  may 
fall  off  as  much  as  10  or  12  per  cent.  The  volume  of  air  passing 
through  the  carburetor  is  equal  (approximately)  to  the  piston 
displacement  of  the  engine  per  unit  of  time  and  may  also  be 
assumed  to  be  constant.  The  weight  of  air,  W,  taken  in  will  then 
be  proportional  to  the  air  density,  D.  At  any  altitude  x, 

W.     Di 

Wa-  D0 

where  o  indicates  ground  condition.  With  a  sharp-edged  fuel 
nozzle  of  constant  coefficient  of  discharge,  the  weight  of  fuel  dis- 
charged, w,  is  proportional  to  the  square  root  of  the  pressure 
drop  at  the  carburetor  throat  (equation  14)  and  this  is,  approxi- 
mately, proportional  to  the  air  density,  D  (equation  12).  This 
may  be  written: 


w 

If  R  is  the  air-fuel  ratio  —  ,  then 

w' 


that  is,  the  strength  of  the  mixture  varies  inversely  as  the  square 
root  of  the  air  density.     As  the  air  density  is  proportional  to  its 


260 


THE  AIRPLANE  ENGINE 


pressure  and  inversely  as  the  absolute  temperature,  T,  this 
becomes 

Ro 

R* 

On  going  from  the  ground  to  an  altitude  of  30,000  ft.,  where  the 
air  density  is  40  per  cent  of  ground  density,  the  air-fuel  ratio 
would  fall  from  20  to  14,  or  the  strength  of  the  mixture  would  be 


enriched 


X  100  =  43  per  cent. 


The  strength  of  mixture  desired  can  only  be  determined  by 
engine  tests.  Such  tests  have  been  carried  out  on  Hispano- 
Suiza  and  Liberty  engines  at  the  Bureau  of  Standards.  l  The  best 


2.4 


o.oi       oxn       o.o5 

Air  Density  in  Lb.  per  Cu.  Ft. 


0.03 


FIG.    188. — Influence  of  air-fuel  ratio  on 
brakejn.e.p.  at  various  air  densities. 


FIG.  189. — Influence  of  air  density  on 
maximum  power  and  maximum  thermal 
efficiency. 


mixture  to  use  depends  on  whether  maximum  power  or  maximum 
economy  is  wanted.  The  curves  of  Fig.  188  show  how  the  brake 
m.e.p.  varies  with  the  mixture  ratio  at  air  densities,  D,  from 
0.075  to  0.025.  As  these  results  are  for  constant  engine  speed, 
they  also  show  the  method  of  variation  of  the  horse  power 
developed.  It  will  be  seen  that  maximum  power,  P,  is  obtained 
with  an  air-fuel  ratio  of  15  at  all  air  densities.  Maximum 
economy  (minimum  fuel  consumption  per  brake  horse-power 
hour)  is  obtained  at-  the  points  crossed  by  the  curve  M;  it  is 
seen  that  at  ground  level  (D  =  0.075)  the  most  economical 
air-fuel  ratio  is  23,  and  that  this  value  diminishes  (richness  in- 
creases) as  the  air  density  decreases.  The  maximum  economy 
curve  runs  very  near  to  the  limit  of  explosibility,  this  limit 
requiring  an  increasingly  rich  mixture  as  the  compression  pressure 

1P.'  S.  TICE:  Nat.  Adv.  Comm.  Aeronautics,  4th  Annual  Report,  p.  624. 


•  THE  CARBURETOR  261 

diminishes.  The  air-fuel  ratio  for  maximum  economy  is  given 
approximately  by 

R  =  106D  +  15 

where  D  is  the  air  density.  In  Fig.  189  the  same  data  are  re- 
plotted  to  show  the  variation  of  brake  mean  effective  pressure,  and 
of  fuel  consumption  per  brake  horse-power  hour,  with  the  air 
density,  both  at  maximum  power,  P,  and  maximum  economy,  M. 

It  is  not  possible  for  a  carburetor  operating  with  wide-open 
throttle  to  give  both  maximum  power  and  maximum  economy 
without  some  kind  of  manual  control  since  the  demands  for  these 
two  conditions  differ  only  in  the  amount  of  fuel  supplied.  The 
condition  of  maximum  economy  alone  is  important,  except  for 
war  purposes.  For  a  flight  of  several  hours'  duration  the  com- 
bined weight  of  an  engine  and  its  fuel  consumption  will  be  less  for 
a  larger  engine  operating  at  maximum  economy  than  for  a 
smaller  engine  operating  at  maximum  power  and  developing  the 
same  total  power.  The  carburetor  should  be  devised  to  give 
maximum  economy  at  full  throttle,  with  a  manual  control  to 
increase  the  fuel  supply  so  as  to  give  maximum  power  if  desired. 

Economy  of  operation  at  low  loads  is  unimportant  in  heavier 
than  air  machines  since  this  condition  of  operation  is  not  possible 
for  other  than  very  short  periods.  In  lighter  than  air  machines 
the  economy  at  low  loads  may  be  of  more  importance.  At  full 
throttle  the  most  economical  air-fuel  ratio  varies  from  23  at  the 
ground  to  19  at  half -ground  density;  for  operation  at  partial 
loads  these  figures  must  be  reduced.  It  is  not  desirable  to 
operate  an  engine  with  the  mixture  giving  maximum  economy 
because  this  mixture  is  «o  close  to  the  limit  of  explosibility  that 
slight  changes  in  condition  might  result  in  exceeding  that  limit. 
Since  the  economy  changes  but  slowly  with  change  of  mixture 
in  the  neighborhood  of  the  optimum  value,  it  is  the  practice 
to  operate  with  smaller  mixture  ratios;  a  value  of  20  at  the  ground 
is  seldom  exceeded. 

The  optimum  mixture  at  partial  loads  may  be  presumed, 
as  at  full  load,  to  be  fairly  near  to  the  upper  explosive  limit. 
This  limit  changes  with  the  load  as  a  result  of  change  in  compres- 
sion pressure  and  temperature  and  of  change  in  the  percentage 
of  diluting  residual  gases  present.  The  compression  pressure 
exerts  considerable  influence  on  the  explosive  properties  of  a 
weak  mixture,  necessitating  the  use  of  a  stronger  mixture  as  the 


262 


THE  AIRPLANE  ENGINE 


load  diminishes.  The  temperature  at  the  end  of  compression  does 
not  change  much  since  the  ratio  of  temperatures  at  the  beginning 
and  end  of  compression  is  a  function  of  the  ratio  of  compression 
which  remains  constant;  it  may  be  presumed  that  the  tempera- 
ture effect  is  negligible.  The- effect  of  charge  dilution  on  ex- 
plosibility  has  been  investigated  for  mixtures  of  air  with  methane 
and  with  natural  gas,  the  diluting  agent  being  C02.1  Some  of  the 
results  of  this  investigation  are  plotted  in  Fig.  190.  It  is  seen 
that  with  20  per  cent  C02,  a  mixture  of  natural  gas  and  air  cannot 
be  made  to  explode  at  atmospheric  pressure  and  temperature; 


246         8        10       12       14       16        18      20      22 
Carbon  Dioxide  *  Per Cenf  by  Volume 

FIG.  190. — Influence  of  carbon  dioxide  dilution  on  the  explosibility  of  a  mixture 
of  natural  gas  and  air. 

as  the  percentage  of  C02  diminishes  the  upper  and  lower  limits 
recede  until  with  no  CO2  present  we  have  the  lower  limit  with 
5.2  per  cent  and  the  upper  with  11.6  per  cent  of  natural  gas 
present.  The  figures  for  a  gasoline-air  mixture  are  probably  not 
very  different.  With  higher  pressures  and  temperature  the 
explosibility  limits  will  be  changed  but  the  method  of  variation 
will  be  the  same. 

The  amount  of  dilution  of  the  charge  by  residual  gases  can 
be  calculated  approximately  if  the  temperature  of  these  gases  is 
assumed.  The  amount  of  such  dilution  will  vary  with  the 
load  since  the  residual  gases  fill  the  clearance  at  exhaust  pressure 
and  are  approximately  constant  in  weight  at  all  loads.  The 
amount  of  such  dilution,  d  =  Wr/Wc  (where  Wr  =  weight  of 
residual  gases  and  W c  =  weight  of  fresh  charge),  is  shown  in 

ELEMENT:  Bureau  of  Mines  Technical  Paper  No.  43;  "The  Influence  of 
Inert  Gases  on  Inflammable  Gaseous  Mixtures." 


THE  CARBURETOR 


263 


Fig.  19 1,1  which  also  shows  the  corresponding  compression 
pressures.  These  curves  are  for  a  ratio  of  compression  of  5.5 
and  must  be  regarded  as  approximations  only.  The  pressure 
at  the  end  of  compression  is  well  above  atmospheric  pressure 
and  as  the  temperature  is  probably  about  1,100°F.  absolute  the 
dilution  can  be  carried  further  than  indicated  in  Fig.  190  without 
exceeding  the  explosive  limit.  The  pressure  and  dilution  of  the 
charge  at  partial  loads  are  such  as  to  demand  a  richer  mixture 


1300' 


0.8  0.6  0.4 

Load  under  Throttle 


0.2 


FIG.   191. — Influence  of  compression  pressure  on  charge  dilution  at  various  air 

densities  and  loads. 

than  at  full  load  if  a  satisfactory  explosion  is  to  be  obtained.  It 
seems  probable  that  the  air-fuel  ratio  for  maximum  economy  does 
not  fall  below  15  for  any  operating  condition  that  is  likely  to  be 
met;  that  is,  the  maximum-economy  mixture  approximates  to 
the  maximum-power  mixture  as  the  air  density  and  load  decrease. 
With  this  in  mind  the  performance  curve  for  carburetors  under 
partial  loads  can  be  examined.  It  would  appear  that  constancy 
of  mixture  ratio  under  varying  load  is  not  desirable,  and  that 
1  P.  S.  TICE,  loc.  cit.,  p.  634. 


264 


THE  AIRPLANE  ENGINE 


a    carburetor    should   show   enrichment   of   the   mixture   with 
diminishing  load. 

Performance  of  Representative  Carburetors. — Several  carbu- 
retors have  been  investigated  at  the  Bureau  of  Standards1  to 

20 


0.08  0.07  0.06  0.05  0.04  0.02 

Air   Density  in  Pounds  per  Cu.  Ft. 

FIG.  192. — Variation  of  air-fuel  ratio  in  Zenith  carburetor. 

ascertain  the  variation  in  air-fuel  ratio  with  variation  (1)  of 
air  density  and  (2)  of  load.  Three  of  these  carburetors  are 
considered  here.  The  Zenith  carburetor,  A,  Fig.  177,  which  is 


0.8  0.6  0.4  0.2 

Load  under  Throttle 

FIG.  193. — Variation  of  air-fuel  ratio  in  Zenith  carburetor. 

described  in  detail  on  page  272,  has  two  jets,  of  which  one  is 
operating  under  constant  discharge  head  to  compensate  for  the 
natural  enrichment  of  the  mixture  with  increase  of  load  which 
would  take  place  if  the  other  or  main  jet  alone  were  used.  The 
i  P.  S.  TICE,  loc.  tit.,  pp.  620-636. 


THE  CARBURETOR 


265 


Stewart- Warner  carburetor,  B,  Fig.  177,  has  the  throttle  in  the 
intake  (anterior)  and  compensates  for  load  changes  by  reducing 


10 


0.08 


0.03 


0.07  0.06  0.05  0.04 

Air    Density  in  Pounds  per  Cu.  Ft. 

FIG.  194. — Variation  of  air-fuel  ratio  in  Stewart- Warner  carburetor. 

the  air  pressure  in  the  float  chamber  as  the  load  increases  by 
means  of  a  passage  connecting  the  choke  discharge  to  the  float 
chamber.  The  Stromberg  carburetor,  C,  is  described  in  detail 

28 


10 


1.0 


0.8  0.6  0.4 

Load    under  Throttle 


FIG.  195. — Variation  of  air-fuel  ratio  in  Stewart- Warner  carburetor. 

on  page  278.     The  results  of  the  investigations  are  exhibited  in 
Figs.  192  to  197.     For  each  carburetor  there  is  shown  the  varia- 


266 


THE  AIRPLANE  ENGINE 


tion  of  air-fuel  ratio  with  constant  throttle  opening  and  variable 
air  density,  and  with  constant  air  density  and  variable  throttle 
opening.  The  absolute  values  of  the  air-fuel  ratio  are  unimpor- 


0.07  0.06  0.05  0.04  0.03 

Air     Density  in  Pounds  per  Cu.  F+. 

FIG.   196. — Variation  of  air-fuel  ratio  in  Stromberg  carburetor. 

tant  in  this  connection  since  they  are  controlled  by  the  size  of  the 
fuel  jet,  which  can  be  readily  changed;  the  method  of  variation 
of  that  ratio  may,  however,  be  considered  as  characteristic  of 
each  type  of  carburetor.  In  the  Zenith  and  Stromberg  carbu- 


18 


12 


B 

8 


0.0758 


D=  0.0498 


D  =  0.030! 


0=0.0704 


1.0 


0.8  0.6  0.4 

Load    under    Thro-Hfle 


0.2 


FIG.   197. — Variation  of  air-fuel  ratio  in  Stromberg  carburetor. 

retors,  the  need  for  an  additional  altitude  control  device  is 
obvious;  the  mixture  ratio  at  full  load  varies  from  19  to  10.5  in 
the  Zenith  (Fig.  192)  and  from  15.5  to  9.5  in  the  Stromberg 


THE  CARBURETOR  267 

(Fig.  196)  as  the  air  density  diminishes  from  0.07  to  0.03.  The 
enrichment  is  considerably  in  excess  of  that  which  has  been  shown 
(Fig.  188)  to  be  necessary.  With  load  variation  at  constant  air 
density,  the  mixture  is  practically  constant  in  the  Zenith  car- 
buretor (Fig.  193),  but  enriches  with  diminution  of  load  in  the 
other  two  (Figs.  195  and  197) ;  it  has  previously  been  shown  (p. 
263)  that  such  enrichment  is  desirable. 

Altimetric  Compensation. — The  importance  of  maintaining 
an  economical  mixture  at  high  altitudes  is  attested  by  general 
experience  in  the  air.  British  tests,  to  ascertain  the  advantages 
of  a  special  altimetric  control  of  the  carburetor,  have  shown  with 
water-cooled  engines  an  increase  in  endurance  from  4  to  4^  hr., 
and  in  ceiling  from  19,000  to  21,000  ft.;  with  air-cooled  cylinders 
an  increase  in  endurance  from  6^2  to  6%  hr.,  of  ceiling  from 
15,000  to  18,000  ft.  and  of  speed  from  84  to  92  miles  per  hour. 
In  addition  to  this  there  is  less  fouling  of  the  spark  plugs,  the 
cylinders  keep  cleaner,  and  there  is  less  danger  of  stalling  the 
engine. 

Viscous  Flow  Carburetor. — It  has  been  shown  (p.  259)  that 
with  a  standard  simple  carburetor  with  sharp-edged  fuel  orifice, 
the  air-fuel  ratio  varies  as  the  square  root  of  the  air  density  with 
full  throttle  and  constant  engine  speed.  There  is  a  possibility 
of  making  this  ratio  constant,  under  varying  air  density,  by  sub- 
stituting for  the  sharp-edged  orifice  a  capillary  passage  in  which 
the  flow  is  entirely  viscous.  The  laws  of  viscous  flow  are  com- 
plicated,1 but,  with  velocities  below  those  of  turbulent  flow,  it  is 
approximately  true  that  the  velocity  of  flow  is  proportional  to 
the  pressure  head.  In  that  case,  referring  to  page  259,  we  have 

W*  =  D*  =  wx 

W0          Do          Wo 
Wx  W0 

or  -  -  =  -— ,  that  is,  the  air-fuel  ratio  remains  constant  with 

Wx  W0' 

varying  air  density. 

Carburetors  have  been  built  embodying  the  above  principle, 
the  viscous  flow  being  obtained  by  the  use  of  long  capillary  tubes, 
or  by  flow  between  flat  discs  or  cones  as  in  Fig.  198.  At  partial 
loads  the  fuel  supply  will  fall  off  in  proportion  to  the  decrease  in 
pressure  head  instead  of  in  proportion  to  the  square  root  of  the 
pressure  head  and  the  mixture  will  consequently  be  too  weak  at 
low  loads.  Load  control  is  obtained  by  raising  or  lowering  the 

1  See  HERSCHEL,  Bureau  of  Mines,  Technologic  Paper  100. 


268 


THE  AIRPLANE  ENGINE 


disc  (or  cone)  of  Fig.  198  and  thereby  changing  the  width  of  the 
capillary  passage;  this  can  be  done  by  interconnection  with  the 
throttle  lever.  The  principal  objection  to  this  type  of  carburetor 
is  that  the  fuel  flow  varies  with  the  fluidity  of  the  oil  and  this 
varies  both  with  the  grade  of  oil  used  and  with  its  temperature 
(Fig.  186).  A  further  difficulty  is  sluggishness  in  response  to 
quick  opening  or  closing  of  the  throttle  valve. 


\ 


FIG.  198. — Diagrams  of  viscous  flow  carburetors. 

Altimetric  Control. — The  only  practical  method  at  present 
available  for  adjusting  the  air-fuel  ratio  to  the  desired  value 
at  all  air  densities,  as  well  as  at  all  throttle  positions,  is  by  the 
use  of  an  additional  or  altimetric  control.  A  carburetor  may  be 
designed  so  as  to  give  correct  mixtures  for  varying  load  or  for 
varying  air  density  but  it  cannot  satisfactorily  meet  both  con- 
ditions, since,  with  the  same  weight  of  air  flowing  the  weight 
of  fuel  will  be  different  in  the  two  cases.  For  example,  the 
weight  flow  of  air  at  half  load  at  the  ground  will  be  the  same  as 
at  full  load  at  an  altitude  where  the  air  has  half  ground  density; 
the  pressure  drop  and  the  fuel  flow  will,  however,  be  different 
in  the  two  cases  and  therefore  the  air-fuel  ratio  will  be  different. 
If  the  carburetor  is  designed  to  give  correct  mixture  at  all  alti- 
tudes at  full  load  there  would  have  to  be  added  to  it  a  load  con- 
trol (preferably  connected  with  the  throttle  valve)  which  would 
enrich  the  mixture  at  partial  loads.  The  other  method  of  pro- 
cedure is,  however,  usual;  the  carburetor  is  designed  to  give  cor- 
rect mixtures  at  full  and  partial  loads,  and  an  altitude  control 
is  installed  to  permit  a  diminution  in  the  fuel  supply  at  higher 
altitudes.  This  control  is  nearly  always  manually  operated 
but  it  can  be  made  automatic  without  much  complication. 

A   diminution    of   fuel   supply  can  be  brought  about  either 


THE  CARBURETOR 


269 


(1)  by  diminishing  the  size  of  the  fuel  orifice,  or  (2)  by  controlling 
the  pressure  head  under  which  the  fuel  is  flowing.  The  former 
is  most  readily  accomplished  by  the  use  of  a  needle  in  the  jet; 
the  latter  is  the  method  generally  employed  because  it  is  less 
sensitive  in  adjustment  and  turns  out  to  be  more  robust  as  a 
structure. 


n 


A  B 

FIG.  199. — Altitude  control  by  regulation  of  the  float-chamber  pressure. 

Schematic  diagrams  of  some  of  the  more  promising  methods 
of  altitude  control  are  shown  in  Figs.  199  and  200. l  For  the 
control  of  the  float-chamber  pressure,  Fig.  199,  the  top  of  the 
float  chamber  must  be  provided  with  a  vent,  a,  to  the  atmosphere, 
and  a  connection,  b,  to  some  place  where  the  pressure  is  less 
than  atmospheric.  The  control  valve  may  be  in  either  of  these 
passages. 


ABC 
FIG.  200. — Altitude  control  by  regulation  of  the  jet  discharge  pressure. 

The  nozzle  outlet  pressure  can  be  controlled  in  several  ways. 
The  position  of  the  outlet  relative  to  the  air  passage  can  be 
changed,  either  by  shifting  the  choke,  as  in  Fig.  200A,  or 
by  shifting  the  outlet.  The  amount  of  air  passing  the  outlet  can 
be  reduced  by  the  use  of  an  auxiliary  air  valve  located  at  a  point 

1  National  Advisory  Committee  for  Aeronautics,  4th  annual  report,  p.  637. 


270  THE  AIRPLANE  ENGINE 

beyond  the  fuel  outlet,  as  in  C.  A  third  method  is  to  admit 
(or  bleed)  air  to  the  fuel  jet  past  the  metering  orifice,  as  in  B, 
thereby  reducing  the  pressure  head  on  the  orifice. 

The  structures  involving  a  small  plug  valve  controlling  an 
air  stream  (Figs.  199  and  2005)  are  the  simplest  and  most  easily 
produced.  Their  regulation  is  comparatively  direct  and  involves 
small  forces  and  a  minimum  of  parts;  furthermore,  they  adapt 
themselves  readily  to  automatic  control.  For  such  reasons,  these 
methods  are  the  ones  usually  encountered  in  service.  The 
objection  to  them  is  that  they  do  not  permit  of  one  setting  for  all 
loads  at  any  given  air  density  but  require  adjustment  for  each 
throttle  position,  if  maximum  economy  is  to  be  maintained. 

The  method  of  Fig.  200A  is  structurally  clumsy  and  would 
complicate  the  carburetor  considerably.  The  method  of  Fig. 
200C,  using  a  balanced  auxiliary  valve,  would  offer  little  resistance 
to  operation  and  little  complication.  Moreover,  the  mixture 
should  be  satisfactory  at  partial  loads  without  further  manipula- 
tion. The  auxiliary  valve  would  have  to  be  large  to  give  com- 
plete compensation  up  to  one-half  ground  density.  A  simple 
calculation  shows  that  for  this  range  the  area  of  the  auxiliary 
port  must  be  approximately  1.5  times  that  of  the  carburetor 
throat. 

Manual  operation  of  the  altitude  control  is  extremely  unde- 
sirable. The  operation  should  be  continuous  as  the  plane 
changes  its  altitude  or  speed  and  can  at  best  be  only  intermittent 
with  manual  operations.  Moreover,  the  pilot  has  no  definite 
means  of  knowing  how  far  to  move  the  control  but  must  rely 
chiefly  on  the  engine  tachometer  readings.  He  can  find  the 
maximum  power  position  but  not  the  more  important  maximum 
economy  position.  As  he  is  already  burdened  with  a  large  num- 
ber of  controls  it  is  much  better  to  make  the  altimetric  compensa- 
tion automatic. 

The  simplest  automatic  operating  device  is  an  aneroid  bellows. 
A  sealed  flexible-walled  chamber  will  expand  under  reduced 
pressure  and  under  increased  temperature,  that  is,  it  will  respond 
to  change  in  air  density.  If  correction  for  pressure  only  is 
desired,  the  bellows  can  contain  a  spring  under  compression  and 
can  be  exhausted  before  sealing  (see  Fig.  212).  Such  devices  can 
only  operate  satisfactorily  if  the  resistance  which  they  have  to 
overcome  is  small  and  if  the  method  of  control  is  such  as  not  to 
disturb  the  compensation  at  partial  loads. 


THE  CARBURETOR  271 

Atomization. — The  preceding  discussion  has  concerned  itself 
with  the  metering  or  mixture-making  characteristics  of  carbure- 
tors. Other  qualities  which  are  of  importance  are  (a)  the  degree 
of  atomization  of  the  fuel  and  the  homogeneity  of  the  mixture; 
(6)  the  pressure  drop  through  the  carburetor  at  wide-open 
throttle;  (c)  satisfactory  idling  performance;  (d)  acceleration. 
All  carburetors,  in  order  to  be  acceptable,  must  be  satisfactory 
not  only  in  mixture  making  but  also  in  these  other  characteristics. 
Favorable  conditions  for  fine  atomization  of  the  fuel  are  high 
velocities  of  the  air  and,  to  a  minor  degree,  of  the  fuel.  The  air 
velocity  is  always  much  greater  than  that  of  the  entering  fuel  and 
the  atomization  is  largely  due  to  the  high  relative  velocity  of  the 
air.  This  is  particularly  marked  if  the  fuel  is  not  discharged  in 
the  axial  direction.  The  use  of  an  anterior  throttle,  as  in  Fig. 
177 B,  by  increasing  the  air  velocity  at  the  jet  improves  atomiza- 
tion at  partial  loads.  The  admission  of  air  before  the  fuel  outlet 
but  past  the  orifice  (see  Fig.  2005)  is  a  further  favorable  condition. 

Good  atomization  may  be  impaired  by  the  impinging  of  the 
mixture  on  obstacles  such  as  a  butterfly  throttle  valve,  placed 
centrally  above  the  jet  (see  Fig.  177 A).  Here  again  an  anterior 
throttle  has  an  advantage.  The  mixture  will  impinge  on  the 
inlet  manifold  and  the  valves  before  getting  into  the  cylinder,  but 
it  is  better  to  have  such  actions  take  place  as  far  away  from  the 
mixing  point  as  possible.  Best  results  have  been  obtained  with 
a  long  pipe  leading  from  the  carburetor  to  the  manifold,  giving 
more  time  for  vaporization  and  the  formation  of  a  homogeneous 
mixture  before  the  mixture  is  taken  into  one  or  other  branch  of 
the  manifold. 

Pressure  drop  through  the  carburetor  has  been  touched  on  in 
page  251  in  the  discussion  of  the  discharge  characteristics  of  the 
air  passage.  Its  importance  is  solely  in  affecting  the  maximum 
power  output. 

Idling. — An  engine  requires  a  richer  mixture  at  lighter  loads. 
When  the  engine  is  cold  a  still  richer  mixture  is  necessary.  None 
of  the  carburetors  in  use  on  airplanes  will  give  a  satisfactory  idling 
mixture  without  the  use  of  some  auxiliary  device.  This  consists  of 
a  fuel  discharge  above  the  throttle  which  utilizes  the  high  vacuum 
above  the  closed  throttle  to  suck  in  the  necessary  amount  of  fuel. 

Acceleration. — It  is  of  importance  that  the  mixture  should 
respond  rapidly  to  sudden  changes  in  load.  If  the  throttle  valve 
is  opened  suddenly,  the  greater  density  and  inertia  of  the  fuel 


272  THE  AIRPLANE  ENGINE 

tend  to  make  the  mixture  too  weak,  with  the  result  that  the 
engine  will  back-fire  or  misfire.  To  avoid  this,  it  is  common  to 
have  an  auxiliary  supply  of  gasoline  which,  at  partial  loads, 
collects  near  the  fuel  outlet  and  is  drawn  on  first  when  the 
throttle  is  suddenly  opened,  keeping  up  the  strength  of  mixture 
until  the  regular  flow  is  established. 

Certain  special  conditions  have  to  be  met  with  by  an  airplane 
carburetor  as  a  result  of  manoeuvres  of  the  plane.  The  changing 
inclination  of  the  plane  will  change  the  hydraulic  head  at  the  jet 
unless  it  is  placed  at  the  center  of  the  float  chamber.  With  the 
usual  non-concentric  arrangement  of  parts  (see  Fig.  177)  it  is 
desirable  to  have  the  float  chamber  placed  in  advance  of  the  jet 
as  this  will  give  a  greater  hydraulic  head  and  richer  mixture  on 
climbing  and  will  cut  down  the  fuel  supply  on  descent  or  diving. 
It  is  necessary  to  see  that  the  gasoline  does  not  overflow  from 
the  jet  when  the  plane  is  resting  on  the  ground.  The  action  of 
the  float  and  float  valves  during  a  dive  must  be  examined.  The 
usual  float,  guided  by  a  central  spindle  which  is  normally  vertical, 
will  go  out  of  action  during  a  dive,  with  the  probable  result  of 
flooding  the  carburetor.  Special  float  mechanisms  are  desirable 
and  have  been  devised.  In  case  of  flooding  during  a  dive,  the 
air  horn  or  intake  pipe  should  be  so  arranged  that  gasoline  can- 
not spill  out  into  the  fuselage.  As  the  air  horn  is  usually  facing 
forward  to  get  the  advantage  of  the  increased  air  pressure  due  to 
the  relative  wind  velocity,  such  spilling  will  occur  unless  the  air 
intake  pipe  is  led  upward  before  being  turned  forward. 

The  usual  dual  carburetor  has  one  float  chamber,  and  one 
air  intake  to  the  two  chokes.  A  dual  air  intake  pipe  is  to  be 
recommended  as  reducing  the  risk  from  back-fire,  by  making  each 
group  of  three  or  four  cylinders  a  separate  unit  so  far  as  carburiza- 
tion  is  concerned.  With  a  common  air  pipe,  back-fire  may  cause 
the  engine  to  stop;  with  double  intake,  back-fire  into  one  intake 
will  not  interfere  with  the  operation  of  the  cylinders  fed  from  the 
other  intake,  the  engine  continues  to  run  and  the  flame  in  the 
back-firing  intake  is  drawn  up  into  the  engine,  reducing  the  risk 
of  fire.  Furthermore,  a  dual  intake  increases  engine  power  by 
diminishing  the  -resistance  to  air  flow. 

CARBURETOR  CONSTRUCTION 

Zenith. — The  carburetor  which  has  been  used  most  for  air- 
plane engines  is  made  by  the  Zenith  Carburetor  Co.  In  this 


THE  CARBURETOR 


273 


carburetor,  an  attempt  is  made  to  maintain  constant  mixture 
strength  at  varying  throttle  positions  by  the  use  of  two  jets  or 
nozzles,  one  of  which,  the  main  jet,  acts  in  the  usual  way,  while 
the  other,  the  compensating  jet.  delivers  an  amount  of  fuel  which 
is  entirely  independent  of  engine  speed  and  load.  This  arrange- 
ment was  devised  by  Baverey  in  1906.  The  main  jet  alone  would 
give  a  mixture  which  is  at  all  times  too  weak,  but  which  becomes 
richer  as  the  engine  speed  and  load  increase;  the  compensating 
jet  alone  would  give  a  mixture  which  is  at  all  times  too  weak  but 
which  becomes  weaker  still  as  the  engine  speed  and  load  increase. 
The  two  jets  working  together  tend  to  compensate  one  another, 
and,  if  properly  proportioned,  will  give  a  mixture  of  fairly  constant 
strength  under  varying  speed  and  load.  This  is  shown  in  Fig. 
193.  In  this  case,  the  jet  sizes  are  No.  140  for  the  main  jet  and 


(«) 

FIG.  201. — Diagram  showing  action  of  the  Zenith  carburetor. 

No.  150  for  the  compensating  jet,  the  number  indicating  the 
cubic  centimeters  of  water  discharged  per  minute  under  a  12-in. 
head.  The  discharge  for  the  compensating  jet  is  under  a  con- 
stant head  of  2  or  3  in.  of  water;  the  main  jet  discharge  is  under 
the  variable  head  due  to  the  pressure  drop  at  the  throat  of  the 
venturi,  which  depends  on  the  size  of  the  throat  and  may  amount 
to  40  in.  of  water  in  usual  designs.  The  arrangement  of  these 
jets  is  shown  diagrammatically  in  Fig.  201,  in  which  a  shows  con- 
ditions at  rest,  and  b  at  full  throttle.  The  main  jet,  G,  is  located 
as  usual;  the  compensating  jet,/,  discharges  into  the  well,  J,  which 
empties  into  a  nozzle,  H,  concentric  with  the  main  jet,  G.  When 
at  rest,  the  levels  in  the  float  chamber,  the  wells,  and  the  nozzles 
G  and  H,  are  the  same.  On  opening  the  throttle,  the  capacity  of 
the  nozzle,  H,  is  so  much  greater  than  that  of  the  jet,  7,  that  the 
well,  J,  is  kept  drained  and  both  air  and  fuel  are  sucked  up  the 

18 


274 


THE  AIRPLANE  ENGINE 


nozzle,  H.  As  the  pressure  in  the  well,  J,  is  atmospheric,  the  dis- 
charge through  /  is  due  to  the  hydrostatic  head  of  the  liquid  in 
the  float  chamber  and  is  therefore  constant.  The  well,  J,  serves 
also  as  an  accelerating  well,  giving  a  body  of  fuel  immediately 
available  on  opening  the  throttle  from  the  idling  position. 
At  low  speed,  when  the  throttle  valve,  T,  is  nearly  closed,  the 
suction  at  the  throat  is  not  sufficient  to  draw  in  any  gasoline 
and  it  enters  only  through  the  idling  device.  This  device,  shown 
diagrammatically  in  Fig.  202a,  consists  of  the  idling  tube,  M, 
within  the  secondary  well,  P,  which  is  inserted  in  the  main  well,  J, 
into  which  the  discharge  from  the  compensating  jet, /,  occurs. 
The  well  P  is  provided  with  a  small  metering  orifice  at  the  bottom 
through  which  gasoline  can  enter  from  J,  and  with  small  air 


FIG.  202. — Diagram  showing  (a)  idling  device  and  (6)  altitude  control  of  the 

Zenith  carburetor. 

holes  at  the  top.  The  idling  tube,  M ,  terminating  opposite  the 
throttle  valve,  is  subjected  to  a  very  strong  suction  whenever  the 
throttle  is  nearly  closed  and  discharges  gasoline  from  the  well  P. 
This  gasoline  meets  the  air  passing  with  great  velocity  through 
the  small  opening  around  the  throttle  valve  and  forms  the  idling 
mixture.  As  the  throttle  is  opened,  the  vacuum  at  the  throttle 
diminishes  while  that  in  the  choke  increases,  so  that  discharge 
through  M  ceases  and  that  through  G  begins. 

The  altitude  control  of  the  Zenith  carburetor  is  shown  diagram- 
matically in  Fig.  2026.  It  is  of  the  type  illustrated  in  Fig.  199A. 
The  float  chamber  is  open  to  the  air  through  screened  air  inlets. 
The  well  J  is  in  open  communication  at  its  top  with  the  float 
chamber.  A  passage,  P,  from  the  float  chamber  to  the  choke 
discharge,  is  fitted  with  a  stop  cock,  L,  which  is  manually  operated 
by  the  pilot.  This  cock  is  closed  at  the  ground  and  is  opened 


THE  CARBURETOR 


275 


gradually  as  higher  altitudes  are  reached;  it  should  be  opened  as 
far  as  is  possible  without  appreciably  diminishing  the  revolutions 
of  the  engine. 

The  actual  construction  of  a  Zenith  carburetor  is  shown  in 
Fig.  203.  Gasoline  enters  the  float  chamber  through  D  and  the 
needle  valve  seat,  S.  As  soon  as  it  reaches  a  predetermined 
height  the  metal  float,  F,  acting  through  the  levers,  B,  and  the 
collar,  Nj  closes  the  needle  valve,  C,  on  its  seat  S.  From  the  float 
chamber  the  gasoline  flows  (1)  through  the  compensating  jet,  /, 


FIG.  203. — Section  of  Zenith  carburetor. 

into  the  bottom  of  the  well,  J,  and  then  through  the  channel,  K, 
to  the  cap  jet,  H,  which  surrounds  the  main  jet,  G,  and  (2)  through 
the  channel,  E,  to  the  main  jet,  G.  The  idling  tube,  M,  is  inside 
the  secondary  well,P,  and  discharges  through  the  passage,  R,  to  an 
opening  (not  shown)  opposite  the  throttle  valve.  The  altitude 
control  valve,  Y,  is  a  tube  which  is  shown  communicating  with 
the  choke  discharge;  the  other  communication  to  the  float 
chamber  is  not  shown.  It  is  operated  by  the  lever  X. 

As  in  other  carburetors,  a  single  float  chamber  is  used  to  supply 
two  air  chokes  if  the  engine  has  six  or  eight  cylinders.     One  air  horn 


276  THE  AIRPLANE  ENGINE 

or  intake  commonly  serves  the  two  chokes  of  a  duplex  carburetor, 
but  it  has  been  found  that  greater  engine  power  can  be  obtained 
if  separate  intakes  are  used.  Tests  of  special  Zenith  carburetors 
for  the  Liberty  engine  showed  maximum  power  developed  with 
separate  air  intakes  about  4  in.  long.1 

The  special  feature  of  the  Zenith  carburetor  which  has  recom- 
mended it  is  the  absence  of  all  moving  parts.  It  is  general 
experience  that  auxiliary  air  valves,  metering  pins,  and  other 
moving  devices  will  stick  at  times  and  cause  irregularity  of 
action.  For  maximum  reliability  and  fool-proofness  the  com- 
pensating device  should  be  fixed. 

The  Claudel  carburetor,  which  has  been  used  very  extensively 
for  airplane  engines,  especially  in  Europe,  is  now  being  made  in 
this  country.  Like  the  Zenith,  the  compensation  for  load  and 
speed  is  made  without  any  moving  parts.  A  general  view  is 
shown  in  Fig.  204.  The  fuel  discharges  into  the  choke  from  a 
diffusor  which  is  shown  assembled  in  Fig.  2056.  The  diffusor 
has  four  concentric  tubes,  the  air  tube  e,  guard  tube  d,  diffusor 
tube  c,  and  idling  tube  a.  The  main  jet  is  in  a  small  plug  screwed 
into  the  bottom  of  the  diffusor.  Air  at  atmospheric  pressure 
enters  the  bottom  of  the  air  tube,  passes  over  the  top  of  the 
guard  tube  (which  prevents  the  fuel  from  overflowing  when  the 
engine  is  at  rest),  then  goes  through  such  holes  in  the  diffusor  as 
are  above  the  fuel  level,  and  out  through  the  nozzle  holes  to  the 
throat  of  the  venturi.  The  fuel  is  at  the  level  shown  when  the 
engine  is  idling  or  at  rest.  As  the  throttle  is  opened,  the  suction 
in  the  diffusor  increases,  thereby  lowering  the  liquid  level  in  the 
diffusor  bore  and  uncovering  progressively  a  series  of  air-bleed 
or  compensating  holes.  Through  these  holes  the  air  rushes  into 
the  ascending  column  of  fuel  and  atomizes  it  as  it  leaves  the 
nozzle  holes  at  the  top.  At  maximum  load  the  diffusor  is  practical- 
ly emptied  and  all  the  air-bleed  holes  are  in  action,  cutting  down 
the  effective  head  on  the  fuel.  The  compensation  is  by  control- 
ling the  jet  outlet  pressure  along  the  lines  indicated  in  Fig. 
2005.  Any  desired  kind  of  compensation  can  be  obtained  by 
appropriate  design  of  the  size  and  location  of  the  air-bleed 
holes. 

The  diffusor  acts  also  as  an  accelerating  well.  When  idling 
the  diffusor  is  out  of  action  and  all  the  fuel  goes  through  the  cen- 

1  Bulletin,  Experimental  Department,  Airplane  Engineering  Division, 
U.  S.  A.,  Jan.,  1919. 


THE  CARBURETOR 


277 


tral  idling  tube,  mixed  with  some  air  entering  compensating  holes 
from  the  air  tube. 


FIG.  204. — Section  of  Claudel  carburetor. 


Holes  Compensating  Hofes 


Guard 

Tube 

0   Oj 

•i 

_zj          AirTub* 

Bore 


Guard  Tube  PJ  1/3 

(d) 
FIG.  205.  —  Details  of  Claudel  diffusor. 


(e) 


The  throttle  is  a  cylindrical  or  barrel  throttle,  bored  out  so  as  to 
form  a  smooth  continuation  of  the  venturi  when  it  is  wide  open. 
It  offers  no  resistance  at  maximum  load  and  consequently  leads 


278 


THE  AIRPLANE  ENGINE 


to  maximum  volumetric  efficiency  and  power.  As  the  idling 
tube  projects  into  the  throttle  space,  the  throttle  is  slotted  out 
wide  enough  to  pass  around  it.  To  diminish  the  area  through 
this  slot  when  the  engine  is  idling  a  screw,  c,  extends  into  the  air 
space.  Advancing  the  screw  lessens  the  air  area  and  enriches 
the  idling  mixture.  Figure  206  shows  the  idling  position. 

Another  feature  of  this  carburetor  is  the  sliding  air  cone,  A 
(Fig.  204),  which  is  controlled  by  an  external  lever.  When  the 
cone  is  raised  to  contact  with  the  venturi,  it  shuts  off  all  air 
supply  and  puts  maximum  suction  on  the  diffusor.  This  greatly 
enriches  the  mixture  and  is  advantageous  for  starting  in  cold 


FIG.    206. — Idling  device  of  the  Claudel 
carburetor. 


FIG.  207. — Section  of  dual  Claudel 
carburetor. 


weather.  The  same  device  is  used  for  altitude  control.  The 
venturi  used  in  airplanes  is  larger  than  is  necessary  at  the 
ground.  At  low  elevations  the  air  cone  is  kept  in  a  raised 
position  in  order  to  increase  the  suction  in  the  diffusor  to  the 
amount  necessary  to  give  the  desired  mixture.  As  elevation  is 
gained  the  air  cone  is  gradually  lowered,  thus  compensating  for 
the  natural  increase  in  richness. 

A  cross-section  through  the  diffusors  and  throttle  valves  of  a 
duplex  Claudel  carburetor  as  used  on  the  Hispano-Suiza  engine 
is  shown  in  Fig.  207. 

The  Stromberg  carburetor,  Fig.  208,  although  structurally 
very  different,  uses  the  same  general  method  of  compensation 


THE  CARBURETOR 


279 


for  speed  and  load  as  the  Claudel.     The  special  features  of  this 
carburetor  are  the  float  mechanism  and  the  double  venturi. 

The  float  (Fig.  208)  is  spherical  or  cylindrical  (with  horizontal 
axis)  and  is  hinged  as  shown  with  the  pivot  toward  the  tail  of  the 


Metering 


Nojj/e 
Air  Horn  Drain 

FIG.  208. — Section  of  Stromberg  carburetor. 

plane.  With  this  mounting,  the  float  is  in  action  during  all 
ordinary  manoeuvres  .of  the  plane  (Fig.  209),  that  is,  it  keeps 
the  needle  valve  closed  with  a  moderate  amount  of  gasoline  in 


FIG.  209. — Diagram  showing  Stromberg  float  chamber  in  different  orientations. 

the  chamber.  If  the  plane  goes  upside  down  the  weight  of  the 
float  will  close  the  valve.  With  the  arrangement  of  Fig.  208 
the  main  jet  will  overflow  into  the  air  inlet  during  a  steep  dive 
with  closed  throttle.  A  duplex  carburetor  arranged  as  in  Fig. 


280 


THE  AIRPLANE  ENGINE 


210,  with  the  float  between  the  two  discharge  jets,  leaves  no 
possibility  of  such  leakage  of  fuel. 


Large  venturi'' 
tube 


Accelerating   well- 
Metering  nozzle- 
Altitude  control  tube'' 
Main  gasoline  channel 


'Float 

----Air  horn  drain  connect/on 


FIG.  210. — Section  of  dual  Stromberg  carburetor. 

The  diagrammatic  sketch  (Fig.  211)  shows  the  metering  jet, 
E,  discharging  into  channel,  A,  with  air-bleed  holes,  D,  through 
which  air  at  atmospheric  pressure  enters  from  the  outer  channel, 
B.  The  outer  channel  is  also  the  accelerating 
well. 

The  fuel  and  the  atomizing  air  are  dis- 
charged radially  into  the  choke  through  a  ring 
of  small  holes,  located  at  the  throat  of  a  small 
venturi  tube.  This  small  venturi  is  concen- 
tric with  a  larger  venturi  and  discharges  at 
its  throat.  The  discharge  pressure  of  the 
small  venturi  is  considerably  below  atmos- 
pheric pressure  and  the  depression  is  still 
greater  at  the  throat  of  the  small  venturi. 
This  results  in  very  high  velocity  for  that  por- 
tion of  the  air  supply  which  passes  through 
the  small  venturi,  giving  good  atomization  of 
the  fuel  without  having  to  make  the  whole 
air  supply  acquire  a  very  high  velocity.  This 
arrangement  gives  a  small  total  pressure  drop  in  the  carburetor, 
and  consequently  high  volumetric  efficiency  of  the  engine. 

The  idling  device  is  a  miniature  carburetor  with  discharge 
just  above  the  closed  throttle.     The  idling  tube  connects  directly 


FIG.  211.  —  Dia- 
gram showing  load 
control  of  Stromberg 
carburetor. 


THE  CARBURETOR 


281 


with  the  main  jet  passage  and  has  a  fuel  nozzle  discharging  into 
a  mixing  chamber  where  it  meets  air  entering  through  holes 
which  are  controlled  by  a  needle  valve.  The  discharge  nozzle 
into  the  main  choke  is  a  slot  of  which  more  is  exposed  as  the 
throttle  moves  from  its  closed  position.  The  increased  opening 
of  the  slot  increases  the  suction  in  the  mixing  chamber,  and 
sucks  up  more  fuel  as  the  throttle  begins  to  open.  With  still 
further  opening  the  suction  at  the  main  discharge  nozzle  increases 
while  that  at  the  idling  nozzle  decreases.  There  is  a  throttle 
position  at  which  fuel  discharges  through  both,  but  with  still 
further  opening  the  idling  nozzle  goes  out  of  action. 


FIG.  212. — Diagram  showing  automatic  altitude  control  attached  to  Stromberg 

carburetor. 

Altitude  compensation  is  effected  by  controlling  the  pressure 
in  the  float  chamber.  An  arrangement  for  automatic  control 
is  shown  diagrammatically  in  Fig.  212.  The  aneroid  chamber, 
A,  which  has  been  exhausted  before  sealing,  is  compressed  by 
the  joint  action  of  the  air  pressure  and  the  spring  B.  As  the  air 
pressure  diminishes  the  aneroid  expands  compressing  the  spring 
and  raising  the  valve  C.  The  valve  point  is  slotted  and  offers  a 
decreasing  aperture  for  the  admission  of  air  as  the  valve  rises. 
Air  is  sucked  through  this  slot  by  the  action  of  the  venturi  at  D, 
and  as  the  only  air  vent  from  the  float  chamber  is  into  the  pipe  E, 
the  pressure  in  the  float  chamber  will  vary  with  position  of  the 
valve  C.  An  additional  manual  control  is  a  necessary  safety 
device. 

The  design  in  Fig.  210  is  especially  adapted  to  a  90-deg.  Vee 
engine  and,  as  previously  pointed  out,  permits  a  position  of  the 


282 


THE  AIRPLANE  ENGINE 


float  chamber  between  the  two  carburetor  outlets  which  largely 
eliminates  the  disturbing  factor  of  changing  inclinations  of  the 
plane.  The  carburetor  barrels  are  water-jacketed  for  high 
altitude  service.  The  main  fuel  nozzles  are  in  an  annular  groove 
around  the  small  venturi.  The  altitude-control  suction  is 
through  the  small  axial  tubes  shown  terminating  at  the  throats 
of  the  small  Venturis  and  consequently  give  the  maximum  possible 
suction  and  range  of  action  of  the  control.  The  altitude  control 
has  a  partial  connection  with  the  throttle  in  such  way  that  the 


FIG.  213. — Sections  of  Miller  carburetor. 

mixture  is  enriched  during  the  latter  part  of  the  closing  of  the 
throttle. 

The  Miller  carburetor  has  been  used  on  the  U.  S.  Bugatti 
engine.  It  is  of  the  multiple-jet  type  in  which  load  compensation 
is  effected  by  bringing  more  jets  into  action  as  the  throttle  is 
opened,  the  sizes  of  the  jets  being  designed  to  give  correct  mix- 
ture at  all  loads.  The  jets  are  air-bled,  giving  compensation 
for  varying  speed.  The  jets  are  held  in  a  narrow  holder  (Fig. 
213)  and  discharge  across  a  diameter  at  the  throat  of  the  venturi. 
The  drill  sizes  for  the  Bugatti  engine  are  No.  76,  which  is  the 
idling  jet,  No.  76,  No.  75,  No.  71,  No.  68,  No.  57,  No.  53.  The 
corresponding  diameters  in  inches  are  0.020,  0.021,  0.026,  0.031, 


THE  CARBURETOR 


283 


0.043,  0.0595;  the  areas  consequently  increase  very  rapidly. 
These  jets  come  into  action  progressively  as  the  throttle  is 
opened.  Each  jet  has  four  small  air  holes  just  above  the  metering 
orifice;  air  enters  at  atmospheric  pressure  through  a  Ke-in- 
hole  near  the  top  of  the  jet  holder  and  passes  down  around  the 
outside  of  each  j  et  to  the  air  holes.  The  gasoline  flows  from  the  float 
chamber  to  the  lower  Ke-in-  h°le  in  the  jet  holder.  The  idling 
jet  is  the  first  in  the  holder. 

The  throttle  valve  is  of  the  barrel  type  bored  out  to  give  a 
venturi  form  when  wide  open.  The  stop  for  the  idling  position 
is  seen  in  the  figure.  Altitude  compensation  is  obtained  by 
varying  the  pressure  in  the  float  chamber,  the  air  space  of  which 
is  at  all  times  in  direct  connection  with  the  venturi.  A  manually- 


6as  Fbssage 
FIG.  214. — Sections  of  Master  carburetor. 

operated  valve  controls  the  size  of  the  free  air  connection  to  the 
top  of  the  float  chamber. 

The  Master  carburetor  is  also  of  the  multiple-jet  type,  but 
differs  from  the  Miller  in  that  the  jets  are  all  of  the  same  size 
and  are  not  air-bled.  The  throttle  is  of  barrel  type  (Fig.  214) 
with  an  opening  that  is  curved  so  as  to  uncover  the  jets  pro- 
gressively as  the  throttle  is  opened.  An  air  damper  controlled 
by  the  pilot  restricts  the  venturi  opening  and  consequently 
enriches  the  mixture  when  desired  for  starting.  The  number 
of  jets  is  usually  from  14  to  21,  which  demands  extremely  small 
metering  orifices. 

The  Ball  and  Ball  carburetor  (Penberthy  Injector  Co.)  is  of 
the  single  metering  orifice,  air-bled  type.  The  float  is  spherical 
in  a  spherical  chamber.  The  venturi  throat,  A,  (Fig.  215),  has 
the  main  nozzle  tubes,  B,  connecting  through  the  annulus,  C, 


284 


THE  AIRPLANE  ENGINE 


with  the  passage,  D,  and  the  mixing  chamber,  E.  The  metering 
jet,  Fj  is  at  the  bottom  of  the  nozzle,  G,  and  the  fuel  overflows 
through  the  four  air  holes,  H,  into  the  chamber,  E,  which  connects 
to  the  outside  air  through  the  passage,  M,  and  the  air  orifice,  N. 
Gasoline  arrives  from  the  float  chamber  at  J.  The  idling  jet,  P, 
connects  through  the  passage,  0,  with  the  mixing  chamber,  E,  and 
discharges  just  above  the  closed  throttle.  An  auxiliary  air 
valve,  S}  is  sometimes  used  to  reduce  the  strength  of  the  mixture 
at  heavy  loads. 


FIG.  215. — Section  of  Ball  and  Ball  carburetor. 

The  altitude  control  is  by  variation  of  the  pressure  on  the  dis- 
charge side  of  the  main  jet.  This  is  accomplished  by  substitut- 
ing a  larger  valve-controlled  opening  for  the  air  orifice,  N ;  opening 
this  valve  increases  the  pressure  on  the  discharge  side  of  the  main 
jet  and  weakens  the  mixture. 

The  carburetor  used  on  the  (German)  Basse-Selve  engine  is 
simpler  and  lighter  than  any  of  the  types  previously  discussed. 
The  float  (Fig.  216)  is  annular,  and  concentric  with  the  choke, 
thereby  reducing  the  possibility  of  overflow  of  gasoline  from  the 
main  jet  when  the  carburetor  is  inclined.  The  main  jet  is 


THE  CARBURETOR 


285 


formed  by  a  hole  drilled  in  a  tube  which  is  screwed  diagonally 
into  the  water-jacketed  body  of  the  carburetor  and  lies  across 
the  choke  tube.  The  jet  tube  is  open  at  its  lower  end  and 
projects  into  the  bottom  of  the  float  chamber.  The  idling  jet 
is  formed  by  a  second  tube  of  small  diameter  inside  the  jet  tube. 
This  idling  tube  is  also  open  at  the  bottom  and  is  drilled  radially 
with  a  small  hole  just  below  the  main  jet.  It  communicates 
with  the  mixing  chamber  just  above  the  throttle  by  a  passage 
drilled  in  the  carburetor  body.  Altitude  compensation  is  by 
varying  the  air  pressure  in  the  float  chamber. 


FIG.  216. — Sections  of  Basse-Selve  carburetor. 

The  float  chamber  is  made  of  pressed  sheet  steel  of  very  light 
gage.  The  needle  valve  (Fig.  216)  is  acted  on  directly  by  the 
float  without  the  intervention  of  levers. 

The  carburetor  of  the  Bayerische  Motoren  Werke  engine  has 
some  noteworthy  features.  It  consists  of  three  carburetors 
with  a  common  float  chamber  (Fig.  217).  Each  of  these  car- 
buretors has  a  separate  discharge  pipe  leading  to  a  common 
induction  manifold.  The  central  carburetor  has  both  idling  and 
main  jets;  the  outer  two  have  main  jets  only.  There  are  five 
throttle  valves  arranged  in  two  systems  with  independent  control. 
The  main  system  has  three  throttles,  one  to  each  carburetor. 
The  secondary  system,  which  is  an  altitude  control,  has  valves  on 
the  outer  carburetor  only. 

The  action  is  as  follows:  When  the  main  throttle  is  opened 
slightly,  the  side  throttles  remaining  closed,  the  idling  jet  (center 
carburetor)  alone  is  in  action;  mixture  from  the  center  carburetor 


286 


THE  AIRPLANE  ENGINE 


alone  reaches  the  cylinders.  As  the  throttle  is  opened  further 
the  main  jet  of  the  center  carburetor  comes  in  action  and  supplies 
the  whole  mixture  until  the  throttle  is  half  open.  After  this, 
the  two  side  carburetors,  which  are  controlled  by  slotted  links, 
begin  to  open.  The  normal  continuous  ground  level  full-power 
operation  is  at  the  point  where  the  side  jets  are  just  about  to 
begin  to  discharge. 

So  long  as  the  secondary  throttles  remain  in  their  closed 
position  with  relatively  small  passages  past  them,  a  compara- 
tively rich  mixture  is  supplied  by  the  side  carburetors.  As 
altitude  is  gained  the  secondary  throttles  are  opened  and  give 
increased  power  while  keeping  the  mixture  of  the  desired  strength. 


Adjusfmerrf- 
ibr  5lon  forming 


Main  Jets 


FIG.  217. — Sections  of  B.M.W.  carburetor. 


An  entirely  different  type  of  carburetor  is  used  on  the  Maybach 
engines  on  large  German  dirigibles.  These  have  been  designed 
to  dispense  with  the  use  of  a  float  chamber  and  to  work  in  con- 
junction with  a  gasoline-pump  system.  The  construction  is 
shown  diagrammatically  in  Fig.  218.  The  throttle  valve,  J,  is  of 
the  rotary-barrel  type  and  admits  carbureted  air  from  N  and 
fresh  air  from  L.  The  throttle  lever  is  interconnected  with  the 
sliding  shutter,  K,  controlling  the  air  that  flows  past  the  jets,  and 
with  a  rotatable  cover,  P,  regulating  the  size  of  the  jets.  Fuel 
from  the  gasoline  pump  enters  an  upper  vessel,  A,  by  the  pipe,  B. 
The  level  in  this  vessel  is  kept  constant  by  an  overflow  pipe,  C, 
which  conducts  the  excess  fuel  back  to  the  supply  tank.  An  air 
vent  fitted  with  a  baffle  plate  is  provided  at  F.  The  fuel  passes 


THE  CARBURETOR 


287 


from  A  through  a  strainer,  M,  to  the  vessel,  D,  whence  it  is 
sucked  through  the  orifice,  H,  into  the  induction  pipe.  Excess 
fuel  in  D  overflows  and  joins  the  excess  from  A  in  the  pipe  C. 
At  the  top  of  vessel  D  two  holes  are  drilled — the  main  and  idling 
jets.  These  orifices  are  controlled  by  the  eccentrically-mounted 
cap,  P,  which  is  rotated  through  interconnection  with  the  throttle 


FIG.  218. — Diagram  of  Maybach  carburetor. 

lever.  The  fuel  has  a  constant  liquid  head  equal  to  the  difference 
in  levels  between  the  liquid  in  A  and  the  level  of  the  orifices;  in 
addition  it  is  subjected  to  the  suction  in  the  passage  above  H. 

In  the  idling  position,  L  is  open  slightly  (Fig.  219),  K  is  closed, 
and  the  idling  jet  only  is  uncovered  by  P.     The  throttle-lever 


Angular  displacement  of  control  lever. 

FIG.  219. — Action  of  the  Maybach  carburetor. 

quadrant  is  marked  with  the  positions  "idling,"  "low  speed," 
"full  power,"  and  "altitude."  As  the  throttle  is  rotated  from 
the  idling  position,  which  demands  a  rich  mixture,  the  shutter  K 
opens  but  the  fuel  opening  does  not  increase  much  till  the  "low 
speed"  position  is  reached;  the  fuel  discharge  increases  in  con- 
sequence both  of  increased  fuel  orifice  and  of  the  increased 


288 


THE  AIRPLANE  ENGINE 


suction  at  H.  The  "full  power"  position  is  not  maximum  power 
but  is  the  maximum  at  which  it  is  desirable  to  operate  the  engine 
at  ground  level.  The  fuel  orifice  is  nearly  wide  open  at  the  full- 
power  position.  With  further  opening  of  the  throttle  the  fresh- 
air  inlet  L  opens  more,  thereby  preventing  the  enrichment  of  the 
mixture  which  otherwise  would  occur  at  high  altitudes  and 
maximum  power. 

It  is  evidently  possible  to  design  the  dimensions  and  the 

interconnections  of  the  three 
orifices  G,  N  and  H  in  such 
way  as  to  give  any  desired 
mixture  to  an  engine  operat- 
ing at  ground  level  and  at 
maximum  power  at  various 
altitudes.  Partial  loads  at 
high  levels  are  not  provided 
for.  This  method  of  meeting 
the  carburetor  problem  is  un- 
desirable because  of  the  com- 
plexity of  the  design  and  the 
practical  impossibility  of  mak- 
ing the  varying  fuel  orifices  of 
the  desired  dimensions.  This 
particular  carburetor  is  very 
heavy  and  offers  a  large  air 
resistance,  thereby  reducing 
the  volumetric  efficiency  and 
power  of  the  engine  which  it 
supplies. 

A  very  simple  type  of  carburetor  is  used  on  the  rotary  Le  Rhone 
engine.  The  air-fuel  mixture  enters  the  rotating  crankcase 
through  a  stationary  hollow  crankshaft.  The  screened  air  supply 
is  controlled  by  a  throttle  which  is  in  the  form  of  a  shutter  (Fig. 
220)  carrying  at  its  lower  end  a  long  metering  pin  which  controls 
the  size  of  the  fuel  jet.  The  pressure  at  which  the  fuel  arrives  at 
the  orifice  is  controlled  by  a  by-pass  valve;  this  serves  to  con- 
trol the  mixture  when  altitude  or  load  is  changed.  The  inherent 
mixture  control  is  irregular  and  uneconomical  with  a  device  of 
this  nature. 


FIG.  220. — Section  of  LeRhone  car- 
buretor. 


CHAPTER  XI 
FUEL  SYSTEMS 

The  following  statement  of  the  requirements  of  the  fuel 
system  of  an  airplane  engine  is  abstracted  from  the  "Handbook 
of  Instructions  for  Airplane  Designers"  prepared  by  the 
Engineering  Division  of  the  U.  S.  Air  Service. 

There  should  always  be  more  than  one  means  of  supplying 
fuel  to  the  engine. 

Main-feed  System. — Gravity  feed  should  be  used  throughout  if  it  is 
possible  to  maintain  a  sufficient  head  with  the  airplane  at  maximum  angles 
of  flight.  It  has  been  found  that  a  head  of  18  to  30  in.  is  required  for 
satisfactory  operation  of  current  types  of  carburetor.  Unless  it  is  possible 
to  maintain  a  sufficient  head  by  gravity,  pumps  must  be  installed  to  supply 
gasoline  from  the  main  tanks  to  the  engines.  Pressure  in  supply  tanks  is  not 
permitted  on  fighting  planes. 

The  main  fuel  pumps  should  have  a  capacity  at  least  50  per  cent  greater 
than  the  maximum  requirement  of  the  engines.  Two  pumps,  other  than 
hand  pumps,  are  desirable,  either  of  which  can  supply  sufficient  fuel.  They 
should  have  automatic  pressure  regulation  to  eliminate  the  use  of  relief 
valves  or  other  means  of  adjusting  the  pressure  at  the  carburetor.  The 
gasoline  pressure  at  the  carburetor  must  always  be  at  least  1  Ib.  and  the 
system  should  be  so  adjusted  that  this  pressure  can  never  rise  above  3  or 
4  Ib.  as  a  result  of  change  in  position  of  the  airplane.  Pumps  capable  of  a 
discharge  pressure  higher  than  4  Ib.  should  have  relief  valves  connected 
between  the  suction  and  discharge,  so  adjusted  as  to  limit  the  maximum 
discharge  pressure  to  4  Ib.  The  fluctuation  of  pressure  at  the  carburetor, 
due  to  pulsations  of  the  pump,  should  not  be  over  25  per  cent.  Where  air 
pressure  is  used,  the  power  air-pump  must  be  capable  of  keeping  a  pressure 
of  2  Ib.  on  the  tanks  at  the  ceiling  of  the  airplane  and  both  spring-  and 
manually-controlled  relief  valves  should  be  furnished,  the  former  set  to 
relieve  at  4  Ib.  per  square  inch. 

Pumps  should  preferably  be  located  below  the  lowest  point  in  the  supply 
system.  If  they  are  located  higher  than  the  bottom  of  the  main  tanks, 
means  must  be  provided  for  admitting  gasoline  from  the  auxiliary  supply  to 
the  suction  side  of  the  pumps.  It  should  be  possible  for  the  pilot  to  make  use 
of  this  connection  during  flight.  A  non-return  valve  must  be  installed  to 
prevent  this  gasoline  from  going  into  the  main  tanks  instead  of  the  pumps. 

Pumps  which  do  not  require  glands  are  preferred,  although  satisfactory 
glands  will  be  accepted ;  in  case  glands  are  used,  they  must  be  so  located  that 
any  leakage  from  the  glands  will  be  drained  to  a  point  outside  of  the  fuselage. 
Pumps  should  preferably  be  connected  to  and  driven  by  the  engine. 

Auxiliary-feed  System. — The  auxiliary-feed  system  supplies  gasoline  to 
the  engine  in  case  of  failure  of  the  main  supply;  this  auxiliary  system  should 
19  289 


290  THE  AIRPLANE  ENGINE 

be  such  that  fuel  can  be  supplied  to  the  engine  in  the  shortest  possible  time 
never  to  exceed  a  period  of  10  sec.  from  the  time  the  pilot  starts  to  make  use 
of  the  auxiliary  system.  For  emergency  use,  gravity  tanks  are  best,  but 
must  not  be  used  if  a  head  of  12  in.  in  level  flight  is  not  obtainable;  they 
should  have  sufficient  capacity  to  operate  the  engines  for  30  min.  at  an 
altitude  of  10,000  ft.  with  wide-open  throttle  and  should  be  so  connected 
to  the  system  that  they  can  be  shut  off  and  used  for  reserve  or  emergency 
only.  They  must  be  so  constructed  or  connected  that  they  can  be  entirely 
emptied  with  the  airplane  inclined  at  maximum  angles  of  flight.  An  over- 
flow pipe  from  the  gravity  tank  returns  any  excess  gasoline  to  one  or  more 
of  the  main  tanks;  this  overflow  must  be  so  constructed  that  there  can  be 
no  gasoline  trapped  in  it  when  the  airplane  is  in  normal  flying  position. 

Unless  there  are  three  means  of  delivering  fuel  to  the  engine,  such  as  two 
engine  or  wind  driven  pumps  and  a  gravity  tank,  a  hand  gasoline  pump  must 
be  provided  which  will  permit  the  pilot,  while  controlling  the  airplane,  to 
pump,  without  undue  exertion,  sufficient  fuel  from  the  main  supply  at  proper 
pressure  for  the  operation  of  all  engines  at  full  throttle.  The  capacity  of  this 
auxiliary  system  must  be  such  that  the  pilot  will  not  need  to  operate  the 
pump  during  more  than  one-third  of  the  time. 

Tanks  should  be  of  tinned  steel  and  of  such  thickness  that  the  tank  will 
stand  5  Ib.  per  square  inch  pressure  on  the  inside  without  undue  distortion. 
Flat  surfaces  are  to  be  avoided.  Wherever  the  width  or  length  (horizon- 
tally) of  a  tank  is  greater  than  12  in.,  a  splash  plate  for  reinforcing  purposes 
must  be  installed  at  least  every  12  in. ;  wherever  the  height  of  a  tank  is  greater 
than  18  in.,  a  splash  plate  for  reinforcing  purposes  must  be  installed  at  least 
every  18  in.  All  seams,  including  the  connection  between  the  splash  plates 
and  the  walls  of  the  tanks,  should  be  riveted  and  soldered.  Copper  or  soft 
iron  rivets  must  be  used  throughout;  the  exposed  parts  of  the  rivets  to  be 
tinned  in  case  iron  rivets  are  used. 

Drains  leading  to  a  point  outside  the  fuselage  must  be  installed  in  the 
bottom  of  each  main  tank.  Fillers  must  be  conveniently  located  on  each 
tank,  and  in  such  a  position  that  t*he  entire  tank  can  be  filled  while  the 
airplane  is  on  the  ground.  A  removable  screen  must  be  installed  at  the  point 
of  filling  of  each  main  tank,  and  also,  if  practicable,  in  the  gravity  tank. 
Vents  must  be  located  at  the  highest  point  on  all  tanks,  usually  in  the  filler 
tube,  except  on  wing  gravity  tanks  where  the  overflow  pipe  shall  act  as  a 
vent. 

Line  and  Carburetor  Strainers. — A  line  strainer,  with  removable  screen 
and  bowl,  must  be  installed  between  the  tanks  and  pumps,  located  as  low  as 
possible  and  in  such  a  position  as  to  be  readily  accessible  for  draining  and 
cleaning.  The  strainer  screen  should  be  of  brass,  bronze  or  copper  of  about 
100-mesh  and  0.005-in.  diameter  wire,  and  should  have  at  least  1  sq.  in. 
for  each  6  gal.  which  must  pass  through  per  hour. 

Each  carburetor  should  be  provided  with  a  strainer  having  a  readily 
removable  screen  of  brass,  bronze  or  copper  of  about  50  mesh  and  approxi- 
mately 0.009-in.  diameter  wire  and  having  an  area  of  at  least  2  sq.  in. 

Service  Pipes  and  Connections. — Service  pipes  should  be  %  in.  outside 
diameter  where  flow  is  30  gal.  per  hour  or  less;  %  in.  outside  diameter  where 
flow  is  between  30  and  60  gal.  per  hour;  and  %  in.  outside  diameter  where 


FUEL  SYSTEMS 


291 


flow  is  between  60  and  100  gal.  per  hour.  All  vent  and  air  tubes  should  be 
Y±  in.  outside  diameter.  Wall  thickness  should  be  0.028  in.  for  ^-in.,  3^2  in. 
for  %-in.,  and  %4  in.  for  ^-in.  and  %-in.  outside  diameter.  All  service 
pipes,  or  tubing,  should  be  seamless  and  of  annealed  copper,  soft  enough  to 
withstand  vibration.  At  all  points  where  the  tubing  is  connected  to  solidly 
mounted  objects,  such  as  pumps  or  tanks,  flexible  connections  must  be 
provided.  The  tubing  must  be  properly  protected  at  points  of  possible 
chafing.  Sharp  bends  are  not  permitted.  Tube  fittings  are  to  be  of  brass 
or  bronze. 


Carburefo. 


r 

Filler  Cap 

-^*A 

mall  compartment 
trf/owinq  into 

I 

rqer  compartment  - 

Main       Tank 

«fe= 

r 

Lock  wire  on  f 

cock  hand fe  •'* 


•Primer 


Valve  with 
spring -fa 
automatic- 
al fycfose 


'Distributing 
valve  acce&ible 
top/i/a 

Strainer 
'  Drain  Cock 
%Drain 


of  Fuselage 
FIG.  221. — Fuel  system  for  a  single-engine  airplane  with  gravity  feed. 


Multi-engine  Installations. — When  more  than  one  engine  is  used,  each 
should  have  its  own  gasoline  system,  consisting  of  pumps,  main  tanks, 
gravity  or  reserve  tank,  distributing  valve  and  other  apparatus  required 
for  a  single  engine  system.  A  cross  connection  with  shut-off  valve  should 
be  provided  so  that  any  engine  can  take  fuel  from  the  tanks  of  the  other 
engines,  and  unless  two  pumping  units,  not  operated  by  hand,  are  provided 
for  each  engine,  a  cross  connection  should  be  provided  so  any  engine  may 
receive  fuel  from  the  pumps  of  the  other  engines. 

Priming  Devices. — A  priming  system  should  be  installed  on  every  engine, 
with  the  priming  pump  mounted  in  the  cockpit  in  an  accessible  posit'on. 


292 


THE  AIRPLANE  ENGINE 


Typical  arrangements  of  the  fuel  system  are  shown  in  Figs. 
221  and  222.  Figure  221  shows  a  system  in  which  the  carburetor 
is  near  the  bottom  of  the  fuselage  so  that  gravity  feed  can  be  em- 
ployed. The  auxiliary  tank  is  most  conveniently  and  simply 
made  a  portion  of  the  main  tank.  A  pump  system  with  auxil- 
iary tank  incorporated  in  one  of  the  wings  is  shown  in  Fig.  222. 

Filler  Cap -.,        (Jravity  Tank 


Primer 


OverflowsightgJass  readi/y 
visible  topi/of 


Pump  priming  valve  fo pilot 
this  priming  connection  may 
be  omitfad  if  pumps  a  re  Mm 
•tanks  and  if  the  gravity  tank 
atalltime  can  provide    \ 
sufficient  head  of  the     ' 
carburetor 


'  Yalve  with  spring  \ 
to  automafr'ca/fy  \ 
close  I 


Lock  Wire 
on  cock  handle 


Distributing  valve   \ 
accessible  to  pi /of  \ 


•g  Drain 


FIG.  222. — Fuel  system  for  a  single-engine  airplane  with  pump  feed. 

Pumps. — A  simple  form  of  air  pump,  used  in  the  Hispano- 
Suiza  engine,  is  shown  in  Fig.  223.  It  is  operated  by  a  cam 
on  the  camshaft  which  gives  the  pump  its  compression  stroke; 
the  return  stroke  is  by  the  action  of  the  spring.  A  cup  leather 
on  the  piston  acts  as  a  suction  valve  on  the  return  stroke.  The 
Mercedes  pump  (Fig.  224),  which  is  driven  from  the  end  of  the 
camshaft,  takes  in  air  through  ports  uncovered  by  the  piston  near 
the  end  of  the  suction  stroke.  A  relief  valve  is  incorporated  in 
the  pump. 

Fuel  pumps  are  made  in  many  forms  and  are  driven  either 
from  the  engine  or  by  small  windmills.  Sliding  vane  and  gear 


FUEL  SYSTEMS 


293 


FIG.  223. — Hispano-Suiza  air  pump. 


To  Tank 


AirlnlefPo. 


FIG.  224. — Mercedes  air  pump. 
t 


FIG.  225. — Maybach  fuel  pump. 


294 


THE  AIRPLANE  ENGINE 


pumps  (see  p.  338)  are  often  used  and  differ  from  the  oil  pumps 
only  in  smaller  capacity.  The  compact  duplex  reciprocating 
pump  of  the  Maybach  engine  (Fig.  225)  is  driven  from  a  crank 
on  the  end  of  the  oil-pump  shaft  through  a  yoke  with  a  sliding 
bushing.  Any  leakage  of  gasoline  past  the  plungers  is  into  the 
crank  chamber,  which  is  filled  with  lubricating  oil  under  pressure. 
Another  method  of  avoiding  the  use  of  glands  past  which  fuel 
leakage  might  occur  is  the  employment  of  castor  oil  as  the  dis- 
placing medium.  In  the  Benz  engine  (Fig.  227)  the  fuel  pump 
is  driven  by  worm  gearing  from  the  end  of  the  inlet  camshaft. 


...--  Pressure  Re  fief  Vafve 


To  Carbure-hors 


Top  of  Gasoline  ••' 
Tank 


Tachometer  Drive 


Gasoline  Delivery  to 

Pressure  ffeservo/r 

in  Main  Tank  and  Hand 

Pump  and  Auxiliary 

Tanks 

—  Outlet  Check  Yalve 


'•  Gasoline  Supply 
from  Main  and 
Auxilfiaru  Tanks 


FIG.  226. — Benz  pressure  reservoir. 


FIG.  227. — Benz  fuel  pump. 


The  lower  portion  of  the  cylinder  is  near  the  bottom  of  a  chamber 
containing  castor  oil  and  the  reciprocation  of  the  piston  produces 
a  rise  and  fall  of  the  castor  oil  in  the  annular  space  around  the 
cylinder.  The  castor  oil  acts  like  an  annular  piston  sucking  in 
gasoline  as  its  level  falls  and  discharging  it  as  the  level  rises. 
As  the  speed  of  the  pump  is  slow  (worm-gear  reduction  10.75  to 
1)  it  is  necessary  to  keep  the  discharged  gasoline  under  air  pres- 
sure during  the  suction  stroke  and  this  is  accomplished  by  the 
use  of  a  pressure  reservoir  (Fig.  226)  located  in  the  main  fuel 
tank;  the  pressure  reservoir  also  serves  to  damp  out  pressure 
pulsations. 


CHAPTER  XII 
IGNITION 

Ignition  is  produced  by  the  passage  of  an  electric  arc  through 
the  explosive  mixture,  at  a  time  which  varies  somewhat  with 
operating  conditions,  but  in  airplane  practice  is  about  25  to  30 
deg.  before  dead  center  on  the  compression  stroke.  At  the 
operating  speed  used  in  airplane  engines  the  moving  electrode 
of  the  make-and-break  system  is  impracticable.  The  spark 
passes  between  stationary  electrodes  and  is  incorporated  in 
"spark  plugs"  which  are  screwed  into  the  cylinder  head.  For 
the  production  of  the  electric  arc  the  following  pieces  of  apparatus 
are  necessary. 

1.  A  source  of  electric  energy;  this  may  be  a  primary  or 
secondary  (storage)  battery,  or  more  usually,  a  magneto. 

2.  As  the  potential  required  to  cause  arcing  is  very  large  the 
low  potential  current  generated  in  a  battery   or   low-tension 
magneto  has  to  be  transformed  into  a  high-potential  current  by 
the  use  of  an  induction  coil;  this  is  usually  incorporated  in  the 
magneto. 

3.  The  current  from  a  single  source  has  to  be  sent  in  succession 
to  each  of  several  cylinders;  this  is  accomplished  by  the  use  of  a 
distributor  which  is  located  in  the  high-tension  circuit. 

4.  The  distributor  connects  up  the  circuit  to  that  cylinder  in 
which  ignition  is  next  to  occur  and  maintains  that  connection 
throughout  a  short  period.     The  actual  timing  of  the  ignition 
within  that  period  is  controlled  by  a  timer,  breaker,  or  interrupter 
located  in  the  low-tension  circuit. 

Other  minor  but  essential  elements  will  be  discussed  later. 

Electric  ignition  systems  utilize  electro-magnetic  phenomena. 
An  electric  current  is  induced  whenever  a  conductor  is  moved 
through  a  magnetic  field  or  when  the  magnetic  field  around  a 
conductor  is  varied.  The  intensity  of  the  induced  current  is 
proportional  to  the  rate  at  which  the  conductor  cuts  the  lines  of 
magnetic  force  and  to  the  number  of  coils  cutting  the  lines  of 
force. 

295 


296  THE  AIRPLANE    ENGINE 

The  simplest  kind  of  electric  ignition  system  is  shown  in  Fig. 
228;  B  is  a  source  of  current,  N1  a  coil  of  wire  surrounding  an  iron 
core  (forming  an  electric  magnet),  S  is  a  switch,  timer,  or  other 
device  for  breaking  the  circuit  at  any  desired  moment.  The 
magnetic  flux  is  represented  by  the  arrowed  lines.  If  the  switch, 
S,  is  opened  the  current  falls  to  zero  and  N'  is  surrounded  by  a 
diminishing  magnetic  field;  if  S  is  closed 
N'  is  surrounded  by  a  rising  field.  In  both 
cases  self-induction  occurs  in  the  coil  Nf 
and  a  current  is  generated  in  it,  whose  mag- 
nitude depends  on  the  rate  at  which  the 
magnetic  field  through  Nf  changes  and  on 
the  number  of  turns  in  the  coil.  The 
FIG.  228. — inductance  phenomena  on  closing  and  on  opening  S 
n  are  quite  different.  On  closing  the  cir- 
cuit, current  can  flow  only  after  the  switch 
is  actually  closed  and  the  flow  is  opposed  by  the  resistance  of 
the  circuit  (which  is  small)  and  the  self-induction  pressure. 
When,  however,  S  is  opened,  an  air  gap  of  great  resistance  is  in- 
troduced into  the  circuit  with  the  result  that  the  current 
diminishes  very  rapidly  and  therefore  establishes  a  high 
electromotive  force  by  self-induction  in  N'.  This  electromotive 
force  is  sufficient  to  overcome  the  resistance  of  the  small  air  gap 
formed  at  the  instant  of  breaking  contact  and  an  arc  is  estab- 
lished across  the  gap.  The  resistance  of  the  arc  is  considerably 
less  than  that  of  the  air  gap  so  that  the  current  may  continue 
to  flow  for  a  short  time  across  a  considerable  arc.  The  more 
rapid  the  opening  of  the  gap,  the  longer  will  be  the  arc.  The 
energy  for  the  arc  is  almost  entirely  the  magnetic  flux  through 
the  coil  N'  and  is  of  comparatively  small  magnitude. 
If  an  additional  or  secondary  coil  N"  be  . 

wound  concentric  with  the  primary  coil  N',    JL+  \  ^]" 


"k 


\" 


as  in  Fig.  229,  the  same  magnetic  changes   ^-B     c> 
will  occur    in  both  coils  and  an  electro- 
motive   force    will    be    generated    in   N"       pIQ    229. High-ten- 

which  is  proportional  to  the  number  of  sion  or  jump-spark  igni- 
turns  in  the  coil.  When  the  number  of 
turns  is  very  large,  a  high  tension,  sufficient  to  jump  an  air  gap 
such  as  ab}  will  be  produced.  A  coil  wound  as  in  Fig.  228  is  called 
an  inductance  or  spark  coil.  A  coil  with  a  double  winding  as 
in  Fig.  229  is  called  an  induction  coil  or  jump-spark  coil. 


IGNITION 


297 


The  formation  of  an  arc  results  in  the  vaporization  or  burn- 
ing of  the  metal  of  one  of  the  points  between  which  the  arc 
springs  and  results  in  deterioration  of  that  point.  To  reduce  the 
arcing  at  S,  a  condenser  is  shunted  around  it.  The  condenser 
consists  of  two  conductors  separated  by  insulating  material; 
it  is  usually  made  of  a  large  number  of  sheets  of  very  thin  metal, 
such  as  tin  foil,  separated  by  thin  paraffined  paper  sheets.  Every 
other  sheet  of  metal  extends  to  one  side  and  the  balance  to  the 
other.  All  the  sheets  of  one  side  are  connected  to  one  terminal 
and  the  remainder  to  another.  By  connecting  the  condenser 
across  the  switch,  S,  (Fig.  229)  the  energy  which  would  otherwise 
go  into  the  formation  of  an  arc  is  absorbed  in  the  system. 


T 


FIG.  230. — Circuit  diagram  of  battery  ignition  system. 

The  schematic  arrangement  of  a  battery  ignition  system  for  a 
four-cylinder  engine  is  shown  in  Fig.  230.  The  primary  circuit 
includes  the  battery,  B,  switch,  S,  primary  winding  on  the  induc- 
tion coil,  I,  the  interrupter,  breaker,  or  timer,  T,  which  breaks  the 
primary  circuit  whenever  ignition  is  required,  and  the  condenser, 
C,  shunted  around  the  timer  to  prevent  arcing.  The  secondary 
circuit  consists  of  the  secondary  winding  of  the  induction  coil, 
the  distributor,  D,  and  the  spark  plugs,  p,p,p,p,;  a  safety  spark 
gap,  G,  (see  p.  310)  is  shunted  on  this  circuit.  The  revolving  arm 
of  the  distributor,  D,  establishes  contacts  successively  with  the 
four  spark  plugs  in  any  desired  order;  the  interrupter,  !T,  breaks 
the  primary  circuit  and  the  current  thereby  generated  in  the 
secondary  circuit  arcs  across  the  spark  plugs.  The  circuits  are 
grounded  as  indicated. 

The  Magneto. — Most  airplane  engines  at  the  present  day  have 
magnetos  as  sources  of  electric  current.  A  magneto  differs 


298 


THE  AIRPLANE  ENGINE 


from  a  dynamo  or  electric  generator  in  having  permanent  mag- 
nets in  place  of  electro-magnets  for  the  fields. 

In  Fig.  231  is  shown  the  action  of  an  armature  type  magneto, 
consisting  of  pole  pieces,  N,  S,  which  are  permanent  magnets, 
and  an  armature,  AB,  consisting  of  core  and  end  pieces,  revolving 
between  the  shoes  of  the  pole  pieces.  The  clearance  ("air  gap") 
between  armature  end  pieces  and  magnet  shoes  is  only  about 
0.005  in.  A  coil  is  wound  on  the  armature  core,  one  end  of  the 
coil  being  grounded;  the  other  end  is  carried  away,  insulated, 
through  a  collector  ring  and  brush.  As  the  armature  revolves 
(being  driven  from  the  engine  shaft)  the  lines  of  magnetic  force 
take  the  successive  directions  indicated  by  the  long  arrows.  The 
magnetic  circuit  is  NABS  for  positions  I  and  II.  In  the  vertical 
position  flux  through  the  core  ceases,  and  no  current  is  generated 


FIG.  231. — Armature  type  magneto. 

in  the  coil.  As  the  armature  passes  the  vertical  position,  the 
circuit  reverses  to  NBAS.  This  continues  for  180  deg.  more, 
when  the  original  direction  of  flow  is  restored.  The  strength  of 
the  magnetic  field  influencing  the  armature  coil  is  greatest  at 
horizontal  positions  of  the  armature;  but  the  rate  of  change  of 
field  strength  is  greater  near  the  vertical  positions,  where  the 
direction  of  magnetic  flux  is  reversing  itself.  The  air  gap  (in  a 
construction  like  Fig.  231)  is  then  large,  so  that  the  maximum 
effective  rate  of  change  occurs  shortly  after  leaving  positions  II 
and  V.  Hence  at  these  positions,  twice  in  every  revolution  of 
the  armature,  the  induced  current  reaches  a  maximum  value, 
and  is  capable  of  producing  a  vigorous  spark. 

Figure  232  shows  the  method  of  variation  of  the  induced 
current  with  magneto  position.  Starting  at  position  II,  Fig.  231, 
the  magnetic  flux  begins  to  diminish  and  has  completely  reversed 
itself  by  the  time  position  III  is  reached.  The  duration  of 
this  period  depends  on  the  width  of  the  armature  end  pieces. 
From  position  III  to  V  there  is  practically  no  induced  current. 


IGNITION 


299 


A  high-tension  magneto  differs  from  that  just  described  in 
that  it  has  both  primary  and  secondary  coils  wound  on  the 
same  armature.  Both  coils  link  with  the  same  magnetic  circuit 
and  therefore  the  armature  becomes  an  induction  coil  and 
replaces  the  separate  induction  coil 
which  would  otherwise  be  necessary. 
Figure  233  shows  a  high-tension  mag- 
neto. The  two  windings  are  shown 
with  one  terminal  grounded  to  the 
machine.  The  primary  coil  is  short- 
circuited  by  the  contact  at  the  inter- 
rupter, at  M,  until  the  proper  moment, 
when  it  is  opened  suddenly  and  the 
induced  high-tension  current  goes 
through  the  distributor  to  one  of  the 
spark  plugs.  The  magneto  and  inter- 
rupter must  be  properly  synchronized 
so  that  the  break  occurs  when  the  primary  e.m.f.  is  a  maximum. 
The  switch,  when  closed,  short-circuits  the  primary  circuit  and 
thereby  prevents  the  building  up  of  a  high-tension  current  in  the 
secondary  circuit,  and  so  shuts  off  the  ignition. 

Of  the  elements  shown  in  Fig.  233  the  condenser  and  inter- 
rupter are  usually  incorporated  in  the  actual  construction  of  the 

Spark  P/ugs 


EL  3E 

Position  of  Armature 

FIG.  232. — Induced  current  in 
armature  type  magneto. 


Condenser . 

"Ground" 
FIG.  233. — Circuit  diagram  of  high-tension  magneto  ignition  system. 

magneto.     The  distributor  may  also  be  incorporated  when  the 
magneto  speed  is  one-half  the  engine  speed. 

The  ordinary  construction  of  a  magneto  with-  revolving  armature 
gives  sparks  at  180-deg.  intervals  corresponding  to  the  positions 


300 


THE  AIRPLANE  ENGINE 


(II  and  V,  Fig.  231)  of  maximum  induced  current.  With  Vee 
type  engines  it  may  be  necessary  to  have  unequal  time  in- 
tervals between  sparks;  for  example,  with  a  two-cylinder  45-deg. 
Vee  engine  the  sparks  instead  of  occurring  at  180-deg.  rota- 
tion of  the  armature  should  occur  alternately  at  157^-deg. 
and  202  J^-deg.  intervals.  This  unequal  interval  can  be  obtained 
in  various  ways.  In  one  of  the  constructions  of  the  Bosch 
Magneto  Co.  the  armature  end-piece  is  cut  away  on  opposite 


I  H  EL 

FIG.  234. — Magneto  with  unequal  firing  intervals  (Bosch). 

sides  of  each  half  of  the  core  so  as  to  increase  the  air  gap  and 
the  tips  of  the  pole  shoes  are  also  cut  away  on  diagonally  opposite 
halves  of  the  two  poles  so  as  to  make  the  positions  of  maximum 
induced  current  (II,  Fig.  231)  come  earlier.  The  construction 
and  operation  are  illustrated  in  Fig.  234.  The  large  air  gap,  B, 
effectively  cuts  off  the  lines  of  force.  Maximum  induced  current 
will  occur  shortly  after  the  armature  has  left  the  trailing  pole 
tips  C  -  D  (position  II)  and  also  after  the  armature  has  left 
the  trailing  pole  tips  E-F  (position  IV).  These  two  positions 


FIG.  235. — Inductor  magneto. 

are  made  less  than  180  deg.  apart  as  a  result  of  cutting  away  tips 
of  the  pole  pieces  at  E  and  F. 

The  magneto  with  revolving  armature  has  to  be  provided 
with  insulated  moving  wires,  collector  rings,  brushes,  and  moving 
contacts  to  convey  the  induced  current  from  the  armature  to 
the  stationary  conductors.  To  avoid  this  complication  a  rotor 
or  inductor  type  of  magneto,  with  stationary  windings,  is  often 
used.  Figure  235  shows  a  construction  with  a  rotating  element, 


IGNITION 


301 


or  inductor,  consisting  of  two  cylindrical  segments  of  soft  iron; 
all  the  rest  of  the  magneto  is  stationary.  The  magnetic  condition 
of  the  armature  core  depends  on  the  position  of  the  inductor. 
In  the  positions  A  and  C  the  segments  form  a  magnetic  bridge 
between  the  magnet  poles  and  the  heads  of  the  armature  core;  in 


90°  180° 

Rotation  of  Inductor  - — >• 
FIG.  236. — Induced  current  in  inductor  magneto. 

these  positions  the  magnetic  flux  is  a  maximum.  In  passing  the 
positions  B  and  D  the  magnetic  lines  are  abruptly  changed  in 
direction  and  a  vigorous  induced  current  is  set  up.  The  reversal 
takes  place  four  times  per  revolution  of  the  inductor  and  succeed- 
ing reversals  give  current  in  opposite  directions.  This  inductor 
magneto  can  give  twice  as  many  ignitions 
per  revolution  and  consequently  has  to  be 
rotated  only  half  as  fast  as  the  rotating 
armature  type  of  magneto.  All  the  elec- 
trical connections  are  stationary.  Typical 
current  curves  are  shown  in  Fig.  236. 

Another  construction  of  inductor  magneto 
is  shown  in  Fig.  237.  The  rotor  is  a  steel 
shaft  carrying  two  laminated  soft-iron  arms 
with  a  space  between  them  which  is  occupied 
by  the  stationary  winding  (not  shown) .  The 
arms  project  on  opposite  sides  of  the  shaft 
and  are  of  such  radius  as  to  give  the  smallest 
practicable  air  gap  between  them  and  the 
pole  shoes.  The  magnetic  flux  in  the 
position  shown  is  from  N  to  R,  then  back 
along  the  shaft  to  the  other  arm  and  so 
to  S.  When  the  inductor  has  rotated  180 
deg.  from  the  position  shown,  the  flux  will  be  from  N  to  the  rear 
arm,  then  forward  along  the  shaft  to  R  and  S.  The  flux  through 
the  shaft  is  reversed  twice  every  revolution  and  induces  a  cur- 
rent in  the  winding  around  the  middle  length  of  the  shaft. 


FIG.  237. — Inductor 
magneto  with  two 
arms. 


302 


THE  AIRPLANE  ENGINE 


The  Dixie  magneto  uses  a  different  type  of  inductor.  The 
rotor,  Fig.  238,  consists  of  two  revolving  wings,  N  and  S,  sepa- 
rated by  a  bronze  center-piece,  B.  Each  wing  is  always  in  contact 
with  one  pole  of  the  magneto  (Fig.  239)  and  consequently  keeps 
the  polarity  of  that  pole.  The  rotor  is  surrounded  by  the  field 
structure,  shown  in  Fig.  240,  which  carries  laminated  pole  exten- 
sions on  which  the  winding  with  its  core  is  mounted.  As  the 


FIG.  238. — Rotating  element  of  Dixie  magneto. 

rotor  revolves  the  direction  of  magnetic  flux  through  the  core 
changes  twice  every  revolution. 

Construction  of  Magnetos. — The  constructive  features  of  a 
Bosch  high-tension  magneto  of  the  rotating  armature  type  are 
shown  in  Fig.  241.  The  armature  rotates  at  engine  speed  and 
gives  two  electrical  impulses  per  revolution.  The  distributor 
is  geared  to  the  contact  breaker  and  rotates  at  half  its  speed. 

The  end  of  the  primary  winding  is  connected  to  the  brass 


FIG.  239. — Diagrammatic  outline 
of  Dixie  magneto. 


FIG.  240. — Field  structure 
of  Dixie  magneto. 


plate,  1.  In  the  center  of  this  plate  is  screwed  the  fastening 
screw,  2,  which  serves,  in  the  first  place,  to  hold  the  contact 
breaker  in  its  position,  and,  in  the  second,  to  conduct  the  primary 
current  to  the  platinum  screw  block,  3,  of  the  contact  breaker. 
Screw,  2,  and  screw  block,  3,  are  insulated  from  the  contact 
breaker  disc,  4,  which  has  metallic  connection  with  the  armature 
core.  The  platinum  screw,  5,  goes  through  the  screw  block,  3. 
Pressed  against  this  platinum  screw  by  means  of  a  spring,  6,  is  the 


IGNITION 


303 


contact-breaker  lever,  7,  which  is  connected  to  the  armature  core 
and  to  the  beginning  of  the  primary  winding.  The  primary 
winding  is  therefore  short-circuited  as  long  as  lever,  7,  is  in 
contact  with  the  platinum  screw,  5.  The  circuit  is  interrupted 


19 


Longitudinal     Sec+ion 


Rear    View 


FIG.  241. — Constructive  features  of  Bosch  high-tension  magneto. 

when  the  lever  is  rocked.     A  condenser,  8,  is  connected  in  parallel 
wi-th  the  gap  thus  formed. 

The  end  of  the  secondary  winding  leads  to  the  slip  ring,  9, 
on  which  slides  a  carbon  brush,  10,  which  is  insulated  from  the 


Primary  Wina/incf 
Secondary  Winding 
Frame 


Con-Fact  Breaker  Disk 
FIG.  242. — Wiring  diagram  for  Bosch  high-tension  magneto  ignition  system. 

magneto  frame  by  means  of  the  carbon  holder,  11.  From  the 
brush,  10,  the  secondary  current  is  conducted  to  the  connecting 
bridge,  12,  fitted  with  a  contact-carbon  brush,  13,  and  through 
the  rotating  distributor  piece,  14,  which  carries  a  distributor 
carbon,  15,  to  the  distributor  disk,  16. 


304 


THE  AIRPLANE  ENGINE 


In  the  distributor  disc,  16,  are  embedded  four  metal  segments, 
17.  During  the  rotation  of  the  distributor  carbon,  15,  the  latter 
makes  contact  with  the  respective  segments,  and  connects  the 
secondary  current  with  one  of  the  contacts. 

The  contact  breaker  is  fitted  into  the  rear  end  of  the  armature 
spindle,  which  is  bored  out  and  provided  with  a  keyway.  The 
short-circuiting  and  interrupting  of  the  primary  circuit  is  effected 
by  means  of  the  contact-breaker  lever,  7,  and  the  fiber  rollers,  19. 
As  long  as  the  lever,  7,  is  pressed  against  the  contact  screw,  5,  the  pri- 
mary circuit  is  short-circuited,  and  the  rocking  of  the  levers  by  the 
fiber  rollers,  19,  breaks  the  primary  circuit;  at  the  same  moment 
ignition  takes  place.  The  distance  between  the  platinum  points, 
when  the  lever  is  lifted  on  the  fiber  rollers,  must  not  exceed  0.5 
mm.  (approximately  J^o  in.).  This  distance  may  be  adjusted  by 
means  of  the  screw,  5. 


FIG.  243. — Bosch  interrupter-distributor. 

Another  type  of  Bosch  interrupter-distributor  is  shown  in 
Fig.  243.  This  is  connected  at  1  to  the  engine  cam  shaft  and 
carries  a  cam,  15,  which  has  as  many  lobes  as  there  are  cylinders 
to  the  engine;  four  lobes  are  shown.  The  interrupter  lever,  16, 
is  held  against  this  cam  by  the  spring,  17,  and  when  the  rubbing 
plate  of  the  lever  is  between  two  lobes  the  platinum  points,  18 
and  19,  are  in  contact;  the  contact  is  interrupted  at  the  passage 
of  each  lobe  and  the  primary  circuit  is  thereby  broken.  The 
distributor  rotor  is  on  the  same  shaft  and  carries  a  rectangular 
brass  tube  in  which  is  located  the  carbon  brush,  11.  This  brush 
sweeps  the  cylinder  cavity  in  the  distributor  body,  7,  and  the  con- 
tacts, 8,  which  are  as  numerous  as  the  cylinders.  The  central 
carbon  brush,  10,  keeps  contact  with  the  rectangular  brass  tube. 
Adjustment  of  the  interrupter  is  by  rotation  of  the  whole  dis- 
tributor through  the  timing  arm,  6. 


IGNITION 


305 


Permanent  magnets  are  of  steel  alloyed  with  5  per  cent  of 
tungsten  or  with  chromium.  All  parts  at  which  sparks  may 
occur  (breaker,  distributor,  safety  gap)  should  be  enclosed  to 
reduce  the  fire  risk. 

The  distributor  speed  is  half  the  engine  speed  on  all  stationary 
four-cycle  engines;  the  distributor  may  either  be  incorporated  in 
the  magneto  or  be  driven  direct  by  the  cam  shaft.  The  Dixie 
distributor  for  an  8-cylinder  engine  is  shown  in  Fig.  244.  The 
rotor  carries  two  carbon  brushes,  which  in  an  8-cylinder  engine 
are  180  deg.  —  45deg.  =  1 35  deg.  apart,  and  in  a  12-cylinder  engine 
are  180  deg.  -  30  deg.  =  150  deg.  apart.  These  brushes  are  not 
in  the  same  plane  of  revolution.  In  the  plane  of  the  outer  brush 


Rotor 


Carbon 
Brushes 


Collector  Brushes 
FIG.  244. — Dixie  distributor  for  8-cylinder  engine. 

are  located  four  metal  segments  embedded  in  the  insulating 
distributor  block;  the  other  four  segments  are  located  immedi- 
ately behind  the  first  four  in  the  plane  of  the  second  carbon 
brush.  Contacts  are  made  each  45  deg.  of  rotation  of  the 
distributor  rotor.  The  collector  brushes  are  in  continuous 
contact  with  the  secondary  circuit  and  with  the  carbon  brushes. 
When  rubbing  contact  is  employed  in  a  distributor  a  deposit 
of  carbon  from  the  brush  will  be  left  on  the  distributor  block 
which  must  be  cleared  off  periodically.  To  avoid  this  a  gap 
distributor  is  sometimes  used  with  a  nickel  point  and  a  small 
air  gap  (from  0.01  to  0.02  in.)  across  which  the  current  arcs. 
The  use  of  a  gap  distributor  has  the  additional  advantage  of 
increasing  the  secondary  voltage  when  the  spark  plug  has  its 
resistance  lowered  either  by  carbon  deposit  or  high  temperature, 
20 


306 


THE  AIRPLANE  ENGINE 


and  thereby  giving  a  spark  under  conditions  in  which  it  would 
otherwise  fail  (see  p.  310). 

The  spark  advance  in  airplane  engines  is  generally  fixed  at 
about  30  deg.  A  slightly  greater  advance  is  desirable  at  high 
altitudes,  but  the  advantage  from  its  use  is  so  slight  and  the 
complication  of  an  additional  control  so  undesirable  that  spark 
control  is  not  used  in  airplane  practice.  Spark  adjustment  is 
obtained  by  adjusting  the  breaker.  A  corresponding  adjustment 
of  the  magneto  is  sometimes  provided.  Figure  245  shows  the 
adjustable  driving  gear  arrangement  of  the  magneto  of  the  King- 
Bugatti  engine.  The  bevel  gear  on  the  magneto  shaft  is  fitted 


Packed  with  Soft' Grease 
FIG.  245. — Adjustable  driving  gear  for  Bugatti  engine  magneto.    . 

on  a  taper  with  a  key.  The  gear  has  eight  key  ways,  so  spaced 
that  the  magneto  timing  may  be  set  within  1J^  deg.  of  any 
desired  position.  The  gear  which  adjusts  the  spark  advance  has 
four  internal  spiral  grooves  sliding  over  splines  on  the  sleeve, 
which  is  keyed  to  the  driving  shaft,  but  may  be  moved  along  the 
shaft  by  a  lever.  Movement  of  this  sleeve  revolves  the  magneto 
driving  gear  with  relation  to  the  shaft-driving  gear. 

Adaptation  to  Engine. — In  four-cycle  engines  ha  vingn  cylinders, 

there  are  ~  sparks  necessary  per  engine  revolution.     If  these  are 

supplied  by  a  magneto  giving  m  sparks  per  magneto  revolution, 

the  ratio  of  magneto  speed  to  engine  speed   is  ^.     Since  m 

v  aries  from  one  to  four  (the  interrupter  may  work  only  once  per 


IGNITION  307 

revolution  even  though  conditions  are  right  twice),  the  speed 
ratio  varies  from  -'  to  -.     The  great  majority  of  magnetos  give 

2         o 

two  sparks  per  revolution,  and  run  at  7  times  engine  speed.     A 

multiple-spark  magneto  runs  at  relatively  low  speed. 

In  an  engine  with  equal  ignition  intervals,  one  magneto  may 
serve  any  number  of  cylinders.  Thus  a  two-spark  magneto  for 
15  cylinders  would  run  at  3%  times  engine  speed:  the  same 
magneto  for  9  cylinders  runs  at  2^  times  engine  speed.  With 
unequal  ignition  intervals,  a  special  magneto  (see  page  300)  or  a 
plurality  of  magnetos  must  be  employed. 

The  cycle  of  operations  of  a  jump-spark  ignition  system1 
can  be  considered  as  consisting  of  five  periods.  For  quan- 
titative values  a  typical  magneto  may  be  considered  to  have  the 
constants  given  in  the  following  table. 

CONSTANTS  OF  TYPICAL  MAGNETO 

Primary  turns  (N\) 160 

Secondary  turns  (Nz) 8,000 

Ratio  of  turns  (n) 50 : 1 

Primary  resistance  (Ri) 0.5  ohm 

Secondary  resistance  CR2) 2 , 500  ohms 

Primary  inductance  (Z/i) 0 . 015  henry 

Mutual  inductance  (M) 0.74  henry 

Secondary  inductance  (L2) 36  henrys 

Primary  condenser  (CO 0.2  microfarad 

Secondary  (distributed)  capacity  (C») 50  micro-microfarads 

Normal  speed  of  operation 2,000  r.p.m. 

Primary  current  at  break  (7b) 4  amperes 

Maximum  current  in  spark 0 . 075  amperes 

Breakdown  voltage  of  gap 5 , 000  volts 

Sustaining  voltage  of  gap 600  volts 

Period  I  includes  the  building  up  of  current  in  the  primary 
winding  as  a  result  of  either  the  impressed  voltage  from  a  battery 
or  the  voltage  generated  by  the  rotation  of  a  magneto  armature. 
During  this  period  the  breaker  or  interrupter  is  closed  and  the 
armature  rotates  from  the  position  of  maximum  flux  to  the 
position  where  the  interrupter  opens.  For  the  typical  magneto 
this  corresponds  to  about  100-deg.  rotation  and  lasts  0.008  sec. 
at  2,000  r.p.m.;  the  current  builds  up  to  4  amperes.  Typical 

1  See  Report  58,  5th  Annual  Report,  Nat.  Adv.  Comm.  Aeronautics. 


308 


THE  AIRPLANE  ENGINE 


curves  for  armature  flux  are  shown  in  Fig.  246,  in  which  A  shows 
the  flux  with  open  primary  circuit  and  is  due  to  the  permanent 
magnetos  only.  B  and  C  give  the  total  flux  under  normal  operat- 
ing conditions  at  500  and  2,000  r.p.m.  respectively. 

Period  II  is  the  very  short  period  (about  0.00002  sec.)  extend- 
ing from  the  opening  of  the  interrupter  to  the  breakdown  of  the 
spark  gap  in  the  engine.  During  this  period,  the  magnetic 
energy  of  the  coil  is  in  part  transferred  into  electrostatic  energy 
and  charges  the  condenser  and  the  capacity  of  the  secondary 
circuit.  The  primary  current  flowing  into  the  condenser  against 


50        0        50       100      150      200     250      300 
Angle  of  Rotation  (Degrees) 

FIG.  246. — Typical  curves  for  armature  flux  of  magneto. 

a  constantly  increasing  e.m.f.  will  decrease  at  a  constantly 
increasing  rate.  The  decrease  in  magnetic  flux  resulting  from 
this  decrease  of  primary  current  generates  an  e.m.f.  in  the  sec- 
ondary windings  which  in  turn  sends  a  charging  current  into  the 
distributed  capacity  of  the  secondary  circuit.  If  the  spark  gap 
were  not  present  this  process  would  continue  until  the  magnetic 
energy  had  been  entirely  converted  into  electrostatic  energy; 
the  maximum  voltage  which  would  be  reached  in  the  typical 
magneto  would  be  about  70,000  volts.  As  a  result  of  loss  of 
energy  due  to  resistances,  eddy  currents  in  the  iron  core,  etc., 


IGNITION 


309 


this  maximum  voltage  is  greatly  reduced.  The  curves  of  Fig. 
247  show  the  rate  of  rise  of  secondary  voltage  as  calculated; 
(A)  with  no  energy  loss  except  that  in  the  resistance  of  the  wind- 
ings, (B)  with  the  usual  eddy  currents.  It  is  assumed  that  the 
spark  gap  does  not  break  down  and  there  is  no  arcing  at  the 
interrupter  points. 

Period  III  is  the  very  short  period  (about  0.00005  sec.)  begin- 
ning at  the  instant  at  which  the  spark  gap  breaks  down  and 
lasting  until  a  steady  arc  is  established.  When  this  gap  has 
broken  down  it  affords  a  conducting  path  into  which  the  charged 
secondary  capacity  discharges.  As  the  secondary  is  now  short- 
circuited  by  the  arc  the  current  increases  rapidly.  The  energy 


40,000 


20        40         60         80         100        120       140 
Time  After  Break'YMiliionthsof  a  Second) 

FIG.  247. — Typical  curves  for  secondary  voltage  of  magneto. 

discharged  into  the  gap  during  this  time  is  about  0.002  joule, 
which  is  just  about  sufficient  to  ignite  the  explosive  mixture 
(see  p.  312). 

Period  IV  extends  from  the  establishment  of  the  gap  to  the 
extinction  of  the  spark.  During  this  time  there  is  a  steady 
discharge  which  lasts  for  a  considerable  time  (0.003  sec.  for  a 
5-mm.  spark  gap  in  air).  The  cessation  of  the  arc  is  due  usually 
to  the  exhaustion  of  the  energy  supply,  although  occasionally 
it  may  be  extinguished  by  the  closing  of  the  interrupter  if  the 
r.p.m.  is  very  high  or  the  spark  gap  very  short. 

Period  V  covers  the  remainder  of  the  cycle  during  which 
time  both  circuits  are  practically  free  from  current. 

Of  these  periods,  II  is  the  most  important,  as  it  determines 
whether  or  not  a  spark  passes  at  all;  the  distributed  capacity 
of  the  secondary  circuit  is  of  great  moment  in  determining  the 
maximum  voltage  in  this  period.  While  5,000  to  6,000  volts 


310 


THE  AIRPLANE  ENGINE 


is  usually  required  to  jump  the  gap,  it  may  be  much  increased 
by  oil  films  on  the  points  and  a  cold  cylinder.  If  the  spark  plug 
is  fouled  with  a  conducting  film  of  carbon  some  of  the  energy 
will  be  drained  by  leakage  during  Period  II. 

A  typical  oscillograph  showing  the  variations  of  the  primary 
and  secondary  circuits  is  given  in  Fig.  248.  The  maximum 
current  delivered  to  the  secondary  circuit  is  usually  from  0.05  to 
0.10  amperes. 

A  safety  spark  gap  is  sometimes  shunted  on  the  secondary 
circuit  (Figs.  230,  241,  and  242)  to  prevent  the  formation  of 

excessive  voltages,  and  the 
consequent  possible  break- 
ing down  of  the  insulation, 
in  case  the  secondary  cir- 
cuit is  open.  This  would 
occur  when  a  spark  plug  is 
being  tested  out  of  the  cyl- 
inder and  is  not  grounded. 


0.  050 


'  Secondary  Current 


4  Amp. 


Primary  Current 
275  Second 


FIG.  248. — Typical  oscillograph  of  primary 
and  secondary  currents  in  magneto. 


The  width  of  the  safety 
gap  is  from  %Q  to  %  in., 
the  higher  value  being  used 
for  high  compression  in  the 
engine. 


The  air  surrounding  the  safety  spark  gap  becomes  ionized 
and  ozone  is  liberated.  If  the  air  is  confined  the  ozone  will 
rust  adjacent  steel  parts  and  slowly  decompose  organic  insulating 
materials.  A  rotary  safety  gap  with  one  electrode  on  the  gear 
driving  the  distributor  and  the  other  integral  with  the  distribut- 
ing metal  electrode  is  used  sometimes;  the  air  is  churned  up  and 
expelled  through  a  suitable  gauze  window. 

A  series  or  subsidiary  spark  gap  is  frequently  used  in  order 
to  maintain  sparking  even  when  the  spark  plugs  are  fouled. 
The  series  gap  is  placed  in  the  connection  between  the  plug 
and  the  magneto.1  Investigations  at  the  Bureau  of  Standards 
show  that  it  is  possible,  by  the  use  of  a  series  gap,  on  an  average 
ignition  system,  to  spark  a  plug  having  a  resistance  lowered  to 
only  4,000  ohms  by  fouling.  At  least  100,000  ohms  insulation 
resistance  is  ordinarily  necessary  at  the  plug  if  a  series  gap  is  not 
used.  For  example,  with  secondary  current  limited  to  0.08 

1  For  elementary  theory  and  results  of  tests  see  Report  57,  bth  Annual 
Report,  Nat.  Adv.  Comm.  Aeronautics. 


IGNITION 


311 


amperes  and  insulation  resistance  of  50,000  ohms  the  maximum 
voltage  across  the  air  gap  is  0.08  X  50,000  =  4,000  volts.  This 
is  not  sufficient;  6,000  volts  is  usually  required. 

The  efficacy  of  the  series  spark  gap  is  well  shown  in  Fig.  249, 
giving  the  results  of  some  tests  by  Young  and  Warren.1  The 
four  resistances  indicated  were  put  in  parallel  with  the  spark  plug 
to  simulate  different  degrees  of  fouling.  With  each  resistance  the 
length  of  the  main  gap  was  varied  while  the  series  gap  was 
kept  constant.  The  curves  show  the  maximum  length  of  the 
main  gap  at  which  sparking  oc- 
curred. With  a  parallel  resistance 
of  58,000  ohms  the  main  gap  could 
be  increased  from  0.9  mm.  to  3.4 
mm.  as  the  series  gap  was  increased 
from  zero  to  0.02  in.  The  voltage 
across  the  main  gap  was  also 
measured  and  was  found  to  in- 
crease from  2,200  to  3,600  by  the 
introduction  of  a  0.02-in.  series  gap, 
when  the  shunted  resistance  was 
112,000  ohms. 

The  series  gap  is  sometimes  in- 
tegral with  the  plug;  sometimes 
at  the  plug  but  not  integral  with 
it;  sometimes  in  the  distributor. 
The  amount  of  the  gap  should  be  variable  to  suit  the  degree 
of  fouling  of  the  plug.  For  maximum  effectiveness  the  gap 
should  be  at,  or  integral  with,  the  spark  plug.  These  desider- 
ata are  conflicting.  It  is  not  practicable  to  adjust  subsidiary 
spark  gaps  at  each  spark  plug,  while  the  engine  is  operating; 
it  is  easily  possible  to  have  a  single  adjustable  spark  gap  at  the 
distributor,  but  it  cannot  be  adjusted  to  suit  the  different  degrees 
of  fouling  in  the  different  cylinders  which  it  serves. 

The  effect  of  temperature  and  pressure  on  sparking  voltage 
has  been  investigated  at  the  Bureau  of  Standards.2  Sparking 
voltage  is  a  linear  function  of  the  density  of  the  gas  and  depends 
on  pressure  and  temperature  only  as  they  affect  the  density. 
For  a  typical  spark  plug  with  0.5  mm.  gap  the  sparking  voltage 


0.03       0.04       0.06 
— v  Length  of  Auxiliary  Spark 
Gap  in  Inches 

FIG.  249. — Effect  of  auxiliary  spark 
gap  on  length  of  main  gap. 


Process  of  Ignition,"  The  Automobile  Engineer,  March,  1920. 
2  See  Report  54,  5th  Annual  Report,  National  Advisory  Committee  on  Aero- 
nautics. 


312 


THE  AIRPLANE  ENGINE 


in  air  varies  from  2,800  volts  at  atmospheric  density  to  9,400 
volts  at  a  density  five  times  as  great.  Other  measurements 
indicate  that  the  sparking  voltage  in  an  explosive  mixture  of 
gasoline  in  air  is  about  10  per  cent  less  than  in  pure  air  and 
that  the  change  in  voltage  is  proportional  to  the  percentage  of 
gasoline  present.  Figure  250  shows  the  observed  sparking 
voltage  for  plugs  with  different  gaps;  No.  1,  1.8  mm.  (0.071 
in.);  No.  2,  1.2  mm.  (0.047  in.);  No.  3,  2.2  mm.  (0.086  in.); 
No.  4,  0.5  mm.  (0.020  in.).  The  voltage  required  for  a  spark 
plug  set  at  0.5  mm.,  in  an  aviation  engine  of  moderate  compres- 
sion, is  about  6,000  volts. 


22 

20 
18 

!" 

5    14 


1234567 
Density  Refer-ed  to  Air  at  1  Atm,  and  273  Decj.  Abs.^ent. 

FIG.  250. — Sparking  voltages  for  plugs  with  different  gaps  at  various  air  densities. 

The  sparking  voltage  is  not  affected  appreciably  by  the 
material  of  the  electrodes  but  is  diminished  by  the  use  of  finer 
points.  The  dimensions  of  the  points  are,  however,  determined 
by  considerations  of  mechanical  strength  and  durability. 

The  " fatness"  of  a  spark  has  no  influence  at  all  on  the  power 
developed  in  an  engine.  If  the  current  is  sufficient  to  charge 
the  plug  and  its  connections  to  the  sparking  potential,  the 
maximum  engine  power  will  be  developed.  The  energy  repre- 
sented by  that  condition  is  usually  about  0.002  joule;  the  energy 
per  spark  varies  from  0.03  joule  in  battery  systems  up  to  0.16 
joule  in  the  more  powerful  magnetos.  The  excess  of  energy 
above  that  necessary  for  ignition  has  no  discernible  effect  on  the 
power  developed. 

Battery  Ignition. — In  a  storage  cell,  electrical  energy  is  stored 
as  chemical  energy  but  returns  to  electrical  energy  when  the  cell 


IGNITION  313 

is  connected  to  supply  an  external  circuit.  The  desired  voltage 
is  obtained  by  connecting  a  sufficient  number  of  cells  in  series, 
forming  a  storage  battery. 

The  chemical  reaction  in  a  lead  cell  may  be  expressed  by  the 
following  equation: 

Charge 

Pb02      +         Pb  +    2HzSOt   =       2PbS04       +      2H20 

Positive  and 
Positive  plate  Negative  plate    Electrolyte    negative  plates    Electrolyte 

Discharge 

Discharge  results  in  the  formation  of  lead  sulphate;  charging 
restores  the  plates  to  their  original  conditions  of  lead  sponge 
(negative)  and  lead  peroxide  (positive). 

The  voltage  of  a  fully  charged  idle  cell  is  2.05  to  2.10  volts. 
Discharge  lowers  the  voltage  in  proportion  to  the  current  flowing. 
Complete  discharge  is  reached  at  1.7  volts,  at  the  normal  dis- 
charge rate  fixed  by  the  manufacturer.  The  capacity  of  a 
battery  is  expressed  in  ampere-hours  at  normal  discharge  rate; 
the  capacity  increases  as  the  discharge  rate  decreases.  The 
maximum  discharge  rate  falls  as  the  temperature  decreases. 

The  acid  or  electrolyte  is  an  aqueous  solution  of  density  1 .255 
resulting  from  the  addition  of  1  part  of  sulphuric  acid  of  sp.  gr. 
1.84  to  4}i  parts  (by  volume)  of  distilled  water.  The  strength 
and  density  of  the  solution  fall  as  discharge  progresses;  when 
the  density  falls  below  1.2  the  cell  needs  recharging. 

Self-sustaining  battery  systems  require  a  generator  for 
recharging.  The  system  may  then  be  regarded  as  a  generator 
system  on  which  the  battery  "  floats."  The  generator  furnishes 
low-tension  direct  current,  and  must  have  a  commutator  and 
brushes.  In  starting,  or  at  low  speed  (up  to  650  r.p.m.),  ignition 
current  comes  from  the  battery.  At  some  definite  speed,  say 
650  r.p.m.  of  the  engine,  the  generator  begins  to  supply  the 
ignition  current.  Its  rate  of  delivery  is  then  considerably  in 
excess  of  that  needed  for  ignition  and  the  surplus  goes  to  the 
battery.  The  recharging  rate  has  a  maximum  value  of  10 
amperes  when  the  battery  is  nearly  discharged.  The  delivery 
voltage  of  the  generator  may  be  automatically  controlled  by  a 
potential  regulator. 

The  outlines  of  the  Liberty  engine  battery  ignition  system 


314 


THE  AIRPLANE  ENGINE 


are  shown  in  Fig.  251.     Figure  252  shows  the  circuit  diagram. 
It  includes  a  low-voltage  generator  in  connection  with  a  storage 


Left  Distributor         Generator 

I     Gen.  Arm.  Gen.  Field 


Gen  Fiefct    Gen.  Arm.       Bcrff. 

FIG.  251. — Liberty  engine  ignition  system. 

battery  of  light  weight  and  small  liquid  content.     Through  a 
switch,  the  current  is  sent  to  two  distributors  which  are  mounted 


Non-inductive 
Resistance 


Reverse  Coil 


•Resistance 

Unit. 

Blades  Separately 
Insulated 


IR  •*  IR 

Left  Distributor  Right  Dis+ri bu-hr 

FIG.  252. — Liberty  engine  circuit  diagram. 

on  the  camshaft  housings  and  are  direct-driven  by  the  camshafts. 
Right-hand  distributors  supply  distributor-end  plugs  and  left- 


IGNITION 


315 


hand  distributors  supply  propeller-end  plugs,  there  being  two 
plugs  to  each  cylinder.  The  entire  system,  exclusive  of  plugs 
and  wiring,  weighs  35  lb.,  and  consists  of  generator,  battery, 
switch,  voltage  regulator,  and  two  distributors. 

The  generator  is  shown  in  Fig.  253.     It  is  four-pole,  shunt- 


in.  high,  and  weighs 


lb.     It 


Brush. 


Armature  •  - 


Filial  Coils. 


Lower  End 
Housing. 


Generator  Armature 
Terminal 


'Commutator 


Upper  End 
Housing 


wound,  4>£  in.  diameter, 
is  mounted  with  its  shaft 
vertical  on  top  of  the 
crankcase  between  the 
two  rows  of  cylinders  at 
the  rear  end.  The  arma- 
ture shaft  extends  down- 
ward into  the  crankcase, 
where  it  is  driven  by  the 
auxiliary  gearing  which 
also  drives  the  vertical 
shaft.  Its  speed  is  1.5 
crankshaft  speed.  The 
end  housings  of  the  gen- 
erator field  frame  are  of 
cast  aluminum.  Gearing 
from  the  armature  shaft 
forms  the  tachometer 
drive.  The  upper  hous- 
ing contains  the  ball 
bearing  for  the  shaft, 
which  is  the  only  bearing 
in  the  generator  proper. 
This  housing  also  sup- 
ports the  four  brush 
holders:  two  positive  brushes, 
grounded  to  the  frame. 

The  ground  side  of  the  field  is  grounded  through  the  voltage 
regulator,  (Fig.  252),  the  generator  voltage  being  determined  by 
the  amount  of  current  flowing  through  the  field,  which  in  turn  is 
controlled  by  the  regulator.  The  armature  has  21  slots  and  is 
wave-wound.  Insulation  between  commutator  segments  is  slotted 
down  ^2  in.  below  the  surface  of  the  copper  bars.  The  shaft  is 
hollow,  and  ground  through  the  bearing.  The  maximum  generator 
voltage  is  10  to  10J^  volts.  A  current  of  5  to  6  amperes  may  be 
carried  without  overheating. 


,  Splined  End 
of  Armature 
Shaft 


FIG.  253. — Liberty  engine  generator. 


insulated,    and    two    negative, 


316  THE  AIRPLANE  ENGINE 

The  voltage  regulator  keeps  the  voltage  constant  at  all  speeds 
above  650  r.p.m.  of  the  engine.  It  weighs  1J£  Ib.  and  is  mounted 
in  a  cast  aluminum  cup  on  the  back  of  the  dash  behind  the  switch. 
It  consists  of  a  soft-iron  core  over  which  a  pivoted  iron  armature 
is  so  mounted  as  to  be  normally  held  away  from  the  core  by  an 
adjustable  tension  spring.  When  so  held,  the  generator  field 
current  passes  through  a  tungsten  contact  point  on  the  armature 
to  the  ground. 

The  core  carries  three  windings  (Fig.  252).  The  voltage 
winding  is  of  fine  wire  leading  from  the  positive  terminal  of  the 
generator  armature  to  the  ground.  The  generator  voltage  is 
impressed  on  this  winding.  Increase  in  this  voltage  increases  the 
core  magnetism  and  opens  the  contact  gap  of  the  regulator 
armature.  This  cuts  off  the  direct  flow  of  field  current  and 
decreases  the  armature  voltage  of  the  generator.  The  reverse 
winding  is  superimposed  on  the  voltage  winding  and  is  also  of  fine 
wire,  wound  in  a  reverse  direction.  The  non-inductive  winding 
consists  of  resistance  wire  wound  so  as  to  produce  no  magnetic 
effect  on  the  core  and  so  as  to  be  itself  free  from  induction  due  to 
changing  core-flux.  These  two  windings  are  connected  in 
parallel  from  the  regulator  armature  contact  point  to  the  ground: 
they  form  a  permanent  high-resistance  ground  for  the  field 
current  when  the  contact  is  open.  The  reverse  winding  rap- 
idly destroys  residual  magnetism  and  enables  the  spring  again 
to  close  the  contacts,  which  in  regular  operation  vibrate 
rapidly. 

The  battery  is  designed  for  light  weight  and  no  leakage.  It 
is  7  by  4  by  5^  in.  and  weighs  10)4  Ib.  It  can  provide  3  amperes 
for  3  hr.,  which  is  sufficient  energy  for  dual  ignition  on  12  cylin- 
ders. It  floats  on  the  line  and  normally  supplies  current  only 
when  the  engine  speed  is  under  650  r.p.m.  The  ammeter  shows 
whether  the  battery  is  charging  or  discharging.  Charging  is 
automatic  at  speeds  above  650  r.p.m.  The  hard-rubber  battery 
jar  has  four  compartments  or  cells,  each  cell  (Fig.  254)  con- 
taining 3+  and  4  —  plates,  burned  to  connecting  straps  and 
separated  by  perforated  rubber  with  wood.  Plates  are  3  by  3 
in.,  and  rest  on  %-in.  bottom  ribs.  Above  the  top  of  each  plate 
is  a  flat  sealing  or  baffle  plate  of  hard  rubber.  The  top  of  each 
cell  is  further  sealed  by  a  rubber  cap  through  which  the  lead 
terminal  posts  extend.  These  also  are  sealed  by  gaskets  or 
by  burning. 


IGNITION 


317 


In  normal  position,  the  electrolyte  completely  fills  the  plate 
compartments.  When  the  battery  is  turned  upside  down,  the 
electrolyte  seeps  through  small  holes  to  the  compartment  which 
is  normally  above  the  plates.  This  compartment  is  of  such  capa- 
city as  to  hold  all  of  the  fluid  in  the  cell,  without  danger  of  over- 
flow at  the  vent  plug. 

The  switch,  Fig.  252,  is  located  on  the  dash  and  weighs  1  Ib. 
It  is  built  on  a  Bakelite  base.  The  circuits  are  controlled  by 


Top  Connector     ,.-~  Venj- PlucfWasfrer-  y-  Positive  Term/no f 
•  /  •  /  Negative  Termtna/ 

VintPtucf  ^        Positive  and 

/  Negative  Plate 
Assembly 


Combined  Wood 
ana?  Rubber 
Separator 


FIG,  254. — Section  of  storage  cell. 

two  aluminum  switch-levers  operating  spring-bronze  contact 
fingers  which  connect  with  contacts  molded  in  the  base.  The  left 
lever  supplies  the  lef.t  distributor,  the  right  lever  the  right  dis- 
tributor. The  engine  is  started  on  one  distributor,  and  both 
levers  are  switched  on  only  when  the  speed  is  above  650  r.p.m. 
(To  start  on  both  distributors  would  require  that  the  battery 
supply  two  sets  of  plugs  and  would  also  waste  battery  current 
through  the  generator.)  Resistance  coils  are  mounted  on  the 
back  of  the  switch  in  series  with  the  distributor  circuits.  These 
prevent  an  excessive  flow  of  current  should  the  switch  be  left  on 
with  the  engine  idle.  The  switch  has  four  external  connections: 


318 


THE  AIRPLANE  ENGINE 


positive  battery,  generator  armature,  and  two  to  distributors. 
Two  12-cylinder  distributors  (Fig.  255)  are  used,  each  supply- 
ing one  plug  in  all  cylinders.  Each  weighs  5J-^  Ib.  and  is  7%  in. 
diameter  by  5%  in.  high.  They  are  mounted  one  on  each  of 
the  overhead  camshafts.  The  transformer  coils  and  breaker 
mechanism  are  incorporated.  The  Bakelite  distributor  head 
forms  a  cover  for  the  breaker  mechanism  and  a  seal  for  the  coil. 
The  moving  contact  is  through  a  soft  carbon  brush  bearing  on 
terminals  molded  in  a  hard-rubber  track. 


Breaker 


Rotor  Arm 


^  Ffo-for  Brush 

Low  Tension 
Lead  to  Sw/Jr-h 
FIG.  255. — Liberty  engine  distributor. 


FIG.  256. — Liberty  engine 
breaker  mechanism. 


The  breaker  mechanism  is  operated  by  a  12-lobed  cam,  having 
lobes  spaced  22^  deg.  and  37^  cleg.  (12-cylinder  engine). 
Tungsten  contact  points  are  used.  Two  main-circuit  breakers 
a  and  6,  Fig.  256,  connected  in  parallel,  are  provided  and  are 
timed  to  operate  simultaneously;  the  duplication  is  a  precaution- 
ary measure.  The  auxiliary-circuit  breaker,  c,  is  provided  to 
prevent  the  production  of  a  spark  when  the  engine  is  " rocked" 
or  turned  backward.  This  auxiliary  breaker  is  connected  in 
parallel  with  the  other  two  through  a  resistance  unit  (Fig.  252) 
which  reduces  the  amount  of  current  flowing  through  it  and  is  so 
timed  that  it  opens  slightly  before  the  other  two  when  the  engine 
is  turned  in  a  forward  direction.  The  opening  of  the  main 


IGNITION  319 

breakers  then  results  in  the  production  of  a  spark.  When  the 
engine  is  turned  in  a  backward  direction  the  two  main  breakers 
open  first  and  no  spark  is  produced  due  to  the  fact  that  the 
current  continues  to  flow  through  the  coil  through  the  auxiliary 
breaker  but  in  diminished  quantity  due  to  the  resistance  unit. 
By  the  time  the  circuit  is  opened  at  the  auxiliary  breaker  the 
intensity  of  the  magnetic  field  of  the  coil  has  weakened  to  such 
an  extent  that  no  spark  is  produced.  The  whole  breaker  mechan- 
ism may  be  revolved  to  advance  or  retard  the  spark. 

Spark  Plugs. — The  conditions  to  which  the  spark  plug  is 
subjected  in  aviation  engines  are  difficult  to  meet.  The  require- 
ments are: 

1.  The  maintenance  of  a  gap  having  a  breakdown  voltage 
of  about  6,000  volts. 

2.  The  maintenance  of  an  insulation  resistance  of  at  least 
100,000  ohms. 

3.  Practically  complete  gas  tightness. 

These  conditions  must  be  maintained  under  pressures  of  500  to 
600  Ib.  per  square  inch  while  immersed  in  a  medium  which  alter- 
nates, 15  cycles  per  second,  in  temperature  between  50°  and  2, 500° 
C.  and  in  an  atmosphere  which  tends  to  deposit  soot,  and  possibly 
oil,  on  the  surface  of  the  insulator.  The  inner  end  of  the  insulator 
and  central  electrode  may  have  an  average  temperature  of  900°C. 
while  the  body  of  the  insulator,  well  up  in  the  shell,  is  in  contact 
with  a  jacket  containing  water  at  70°C.  For  successful  operation 
the  insulating  surface  must  remain  clean,  the  insulator  must  not 
fracture  or  disintegrate  under  the  varying  temperatures  and  no 
part  of  the  plug  must  become  hot  enough  to  cause  preignitions 
of  the  charge. 

The  shell  is  of  steel  with  standard  thread.  The  S.  A.  E.  stand- 
ard plug  dimensions  are:  outside  diameter,  18.2  mm.;  pitch 
diameter,  17.22  mm.  ±0.02  mm.;  root  diameter,  16.09  mm.;  16.9 
threads  per  inch;  1.5  mm.  pitch.  In  order  to  keep  down  its 
temperature  it  is  often  made  with  fins  for  radiating  heat.  Investi- 
gations at  the  Bureau  of  Standards1  have  shown  that  brass  shells 
average  from  90°  to  270°F.  hotter  than  steel  shells. 

The  most  common  insulating  materials  are  mica  and  porcelain. 
Other  materials  used  are  fused  quartz,  steatite  and  molded  mater- 
ials such  as  Bakelite  and  Condensite.  Porcelain  of  the  highest 
grade  is  an  excellent  material  except  for  its  brittleness  and  conse- 

1  Report  52,  5th  Annual  Report,  Nat.  Adv.  Comm.  for  Aeronautics. 


320  THE  AIRPLANE  ENGINE 

quent  liability  to  fracture  either  from  temperature  or  mechanical 
effects.  Mica,  while  free  from  this  trouble,  is  more  likely  to  foul 
in  consequence  of  its  rough  surface.  Fused  quartz  is  free  from 
both  the  above  objections. 

The  insulation  resistance  of  these  materials  diminishes  with  rise 
in  temperature.  At  a  temperature  of  900°F.  the  order  of  merit 
of  the  insulating  materials  is  mica,  quartz,  steatite,  porcelain. 
Porcelain  plugs  which  show  a  resistance  of  millions  of  ohms 
when  cold  may  have  a  resistance  of  only  100JOOO  ohms  at  900°F. 

The  most  common  source  of  failure  in  spark  plugs  is  fouling. l 
It  causes  more  than  50  per  cent  of  spark  plug  troubles  and  is 
most  serious  at  high  altitudes.  Fouling  is  due  to  the  deposit  of  a 
layer  of  carbon  and  causes  a  short  circuit.  The  carbon  deposit 
results  from  either  or  both  of  two  causes:  (a)  The  chilling  of 
the  flame  by  a  cool  portion  of  the  plug  and  the  consequent 
incomplete  combustion;  this  effect  is  particularly  common  when 
the  mixture  is  overrich  and  is  of  frequent  occurrence  at  high 
altitudes  with  an  imperfectly  compensated  carburetor,  (b) 
The  decomposition  of  lubricating  oil  which  is  splashed  on  heated 
portions  of  the  insulator.  The  oil  itself  is  an  insulator  and 
when  it  wets  a  layer  of  soot  in  the  plug  it  makes  it  an  insulating 
layer.  Such  a  deposit  chars  and  becomes  more  and  more  con- 
ducting. The  oil  acts  as  a  binding  material  and  also  increases 
the  rate  of  deposition  of  soot  since  the  carbon  particles  in  the 
flame  adhere  to  it  readily.  The  conduction  through  the  deposit 
seems  to  take  place  through  a  narrow  path  where  the  oil  film 
between  the  particles  has  been  broken  down  by  electric  stress. 
Such  fouling  causes  misfiring. 

A  method  of  attempting  to  reduce  the  deposit  of  carbon  is  to 
keep  the  insulation  at  such  high  temperature  that  all  carbon 
deposit  is  burned  off.  This  can  be  accomplished  by  making  the 
insulator  with  petticoats,  ridges  or  other  projections,  but  there 
results  the  danger  of  preignition,  particularly  in  high  compression 
engines  or  in  those  that  are  not  well  cooled.  Another  method  is 
to  shield  the  insulator  with  a  metal  baffle  plate  which  protects 
it  from  oil  spray,  but  if  this  is  done  the  flame  does  not  get  good 
access  to  the  insulator  and  any  deposit  which  has  formed  will 
have  very  little  opportunity  of  being  burned  away.  The  use 
of  a  series  gap  (p.  310)  is  useful  in  maintaining  firing  after  the 
plug  is  fouled. 

1  Report  52,  5th  Annual  Report,  Nat.  Adv.  Comm.  for  Aeronautics. 


IGNITION  321 

Fouling  with  oil,  either  in  the  form  of  a  surface  film  over  the 
electrodes  or  as  a  drop  between  the  points,  will  often  prevent 
firing.  The  breakdown  strength  of  oil  is  several  times  that  of 
air,  and  the  voltage  required  may  easily  exceed  that  which  the 
ignition  system  is  capable  of  delivering.  The  trouble  is  intensi- 
fied if  the  insulation  of  the  plug  is  at  the  same  time  reduced  by  a 
layer  of  soot,  thereby  diminishing  the  maximum  voltage  which 
the  system  can  develop. 

The  oil  trouble  usually  occurs  on  starting  but  may  also  be 
met  when  the  plane  is  recovering  from  a  long  glide  during  which 
the  engine  is  turning  over  slowly  and  pumping  oil  into  the 
cylinders.  It  may  sometimes  be  identified  by  the  sparking  of 
the  safety  gap  (see  p.  310).  It  seems  to  occur  most  often  when 
the  form  of  the  electrodes  is  such  as  to  drain  the  oil  away  by 
capillary  forces.  The  only  real  remedy  is  to  keep  down  the 
amount  of  lubricating  oil  going  to  the  cylinders  at  starting  and 
during  glides. 

Cracking  the  insulator  is  one  of  the  most  common  causes  of 
spark  plug  failure.  The  thickness  of  the  insulator  is  usually 
so  great  that  a  clear  crack  may  not  interfere  with  ignition,  but 
after  a  while  the  cracks  become  filled  with  carbon  and  form  a 
conducting  path. 

Cracking  may  result  from  several  causes.  The  high  temper- 
ature gradient  from  the  hotter  inner  end  of  the  insulator  to  the 
relatively  cold  shell  and  the  consequent  unequal  expansion  is  a 
frequent  cause.  Such  cracks  are  most  likely  to  occur  at  a 
shoulder  or  other  place  where  there  is  a  sudden  change  in  diam- 
eter. Cracking  may  also  occur  if  the  metal  parts  of  the  plug  are 
so  arranged  that  their  relatively  greater  expansion  produces 
pressure  on  the  insulator.  The  mechanical  vibration  of  the 
engine  as  a  whole  may  break  the  porcelain;  such  breakage  often 
occurs  in  the  outer  portion  of  the  porcelain.  There  is  also 
considerable  breakage  from  accidental  mechanical  injury  such  as 
striking  the  plug  with  a  wrench. 

Mica  plugs  are  free  from  this  trouble.  If  porcelain  is  used  it 
should  combine  high  mechanical  strength,  low  modulus  of 
elasticity,  low  coefficient  of  thermal  expansion  and  high  thermal 
conductivity.  The  porcelain  may  also  be  made  in  two  or  more 
pieces-permitting  the  innermost  porcelain  to  heat  and  expand 
considerably,  while  the  outer  pieces  are  cooler.  The  passage  of 
a  spark  through  the  joint  between  the  pieces  is  prevented  by  a 
21 


322 


THE  AIRPLANE  ENGINE 


wrapping  of  mica  around  both  the  shell  and  the  electrode.  It  is 
very  difficult  to  make  such  plugs  gas-tight. 

In  plugs  in  which  the  central  electrode  is  cemented  in  the 
porcelain,  the  differential  expansion  of  the  two  can  be  taken 
care  of  only  if  the  electrode  is  kept  of  small  diameter. 

Breakage  by  mechanical  vibration  can  be  reduced  if  the 
insulation  is  cushioned  by  a  considerable  thickness  of  asbestos 
or  other  packing  material  between  the  shoulder  of  the  insulator 
and  the  bushing.  One-piece  plugs  in  which  the  edge  of  the  shell 
is  crimped  over  the  shoulder  of  the  insulator  are  especially  liable 
to  cracking  at  the  edge  of  the  shell. 


(b)  (c)  (d) 

FIG.  257. — Types  of  spark  plug  construction. 

A  minor  cause  of  failure  is  the  change  in  the  width  of  the 
spark  gap  either  through  warping  of  the  wires  or  corrosion. 
Warping  occurs  only  when  the  wires  are  relatively  long.  Cor- 
rosion is  very  slow  with  the  alloy  commonly  used  (Ni  97  per 
cent,  Mn  3  per  cent),  although  a  chemical  reaction  between  the 
cement  and  the  metal  of  the  electrode  may  sometimes  cause 
the  tip  to  drop  off. 

Gas  leakage  is  an  evil  in  that  it  causes  a  rapid  heating  of  the 
plug  if  its  amount  is  considerable;  such  leakage  is  usually  a 
matter  of  workmanship  rather  than  of  design  of  the  plug.  There 
are  two  joints  to  keep  tight,  that  between  the  central  electrode 
and  the  insulator,  and  that  between  the  insulator  and  the  shell. 

The  general  methods  of  construction  are  shown  in  Fig.  257. 
In  the  screw  bushing,  a,  the  insulator  has  a  shoulder,  one  side 
of  which  is  seated  on  a  shoulder  in  the  shell  while  a  bushing  is 
screwed  down  inside  the  shell  on  the  opposite  side.  A  gasket 


IGNITION 


323 


of  brass,  copper,  asbestos,  or  some  soft  heat-resisting  material  is 
used  and  can  be  placed  on  either  side  of  the  shoulder  of  the 
insulator,  or  on  both.  With  a  mica  insulator  it  is  possible  to 
dispense  with  the  gasket.  To  relieve  the  insulator  from  the 
mechanical  strain  resulting  from  differential  expansion  of  the 
shell  and  the  insulator  such  constructions  as  those  shown 
diagrammatically  in  Fig.  258  may  be  used.  In  a  the  shell  A  and 
sleeve  B  are  made  of  different  metals  and  B  is  of  such  length  as 
to  maintain  constant  pressure  on  the  washer.  The  expansion 
of  the  central  electrode  is  compensated  in  a  similar  manner. 
In  b  the  steel  clamping  nut  is  very  thin  and  flexible.  Expansion 
of  the  central  electrode  is  provided  for  by  a  strong  spring  washer 
under  the  nut  B. 


A     r-S 


(a)  (6) 

FIG.  258. — Special  spark  plug  constructions. 

The  crimped  shell,  b  (Fig.  257),  is  most  common  (Champion, 
Titan,  etc.)  and  is  formed  by  forcing  the  top  edge  of  the  shell 
over  a  gasket,  which  rests  on  the  upper  side  of  the  shoulder  of 
the  insulator.  These  plugs  cannot  be  disassembled. 

The  taper  fit,  c  (Splitdorf),  is  used  with  a  mica  or  steatite 
insulator.  The  mica  does  not  stand  well  the  pressure  exerted 
on  it  during  assembly.  If  a  thin  steel  jacket  is  placed  over  the 
taper  it  protects  the  mica  and  is  flexible  enough  to  form  a  gas- 
tight  fit  with  the  shell. 

A  molded-in  insulator,  d  (Anderson),  consists  of  glass  which  has 
been  forced  between  the  central  electrode  and  the  shell  while  in 
the  molten  state.  It  adheres  to  both  electrode  and  shell  and  is 
gas-tight. 

The  shape  and  arrangement  of  the  electrodes  seem  to  have 
little  effect  on  the  operation  with  the  exceptions  already  noted 
of  the  greater  liability  to  fouling  with  oil  of  plugs  in  which 


324  THE  AIRPLANE  ENGINE 

the  side  wall  of  the  shell  forms  one  of  the  electrodes.  The 
variation  in  breakdown  voltage  with  the  shape  of  the  tips  is 
slight.  With  fine  wires  any  oil  film  at  starting  is  burned  off 
rapidly  but  there  is  greater  liability  to  preignition.  With  a 
central  electrode  consisting  of  a  disc  (Fig.  257 c)  the  danger  of 
short-circuiting  with  carbon  is  increased  while  the  likelihood 
of  complete  fouling  with  oil  is  diminished  and  greater  protection 
is  afforded  to  the  insulating  material  back  of  it. 

The  location  and  number  of  spark  plugs  used  are  important 
in  their  effect  on  engine  performance.  An  effort  should  be  made 
to  reduce  as  much  as  possible  the  distance  through  which  the 
flame  has  to  be  propagated.  The  greater  that  distance,  the 
greater  is  the  liability  to  detonation.  If  the  distance  is  shortened, 
a  higher  compression  may  be  employed  without  detonation. 
Where  a  single  plug  is  used  its  location  should  be  in  the  center 
of  the  head.  Multiple  spark  plugs  are  desirable,  not  only  to 
insure  ignition  in  case  of  failure  of  one  plug  but  also  because  they 
permit  higher  compression  pressures.  With  a  constant  com- 
pression pressure  the  power  is  increased  by  increasing  the  number 
of  plugs;  for  example,  in  a  5J^  by  6^ -in.  four-valve  single- 
cylinder  test  engine  with  a  compression  ratio  of  5.4  and  at 
1,800  r.p.m.,  the  brake  horse  power  was  increased  5.7  per  cent 
by  the  use  of  two  spark  plugs  and  11.1  per  cent  by  the  use  of 
four  plugs. 

A  typical  wiring  diagram  is  shown  in  Fig.  259,  which  shows 
the  wiring  of  a  12-cylinder  Vee  engine  equipped  with  two  mag- 
netos for  regular  operation,  and  a  starting  magneto.  The 
regular  magnetos  also  have  radio  connection. 

The  relative  advantages  of  battery  and  magneto  ignition 
have  been  much  debated.  The  performance  of  the  engine  is 
not  affected  by  the  source  of  primary  current.  With  battery 
ignition  the  engine  is  started  by  turning  it  with  the  current  on 
but  fully  retarded.  With  magneto  ignition  the  engine  is  pulled 
over  a  few  turns  with  no  current  on  so  as  to  fill  the  cylinder  with 
an  explosive  mixture,  and  a  starting  magneto  is  then  operated 
by  hand,  giving  a  shower  of  sparks  in  one  cylinder;  the  starting 
magneto  is  arranged  to  be  fully  retarded.  The  element  of 
danger  in  starting  the  engine  is  eliminated  by  this  latter  method 
of-  operation.  The  battery  requires  more  attention  than  the 
magneto,  particularly  if  the  engine  is  to  stand  idle  for  a  while; 
if  it  runs  down  there  remains  no  means  for  starting  the  engine. 


IGNITION 


325 


The  battery  system  is  also  more  complicated  than  the  magneto 
system  but  it  lends  itself  better  to  irregular  explosion  intervals; 
the  magneto  must  be  of  special  type  in  this  case  (see  p.  300). 
The  generator  in  the  battery  system  requires  attention  to  keep 
commutator  and  brushes  in  good  condition;  there  is  no  corre- 
sponding attention  necessary  with  the  magneto.  The  total 
weight  of  the  magneto  system,  including  dual  magnetos  and  a 
starting  magneto,  is  somewhat  greater  than  that  of  a  battery 


Right  Distributor- 
Left  Distributor. 


PropelterEncf 


'Right  Distributor 
Left  Distributor 


Left  Magneto . 


Radio        .^ 

Connect/on  -*. 


Radio  Connection 


Righf  Magneto 


Left  Distri  butor 
Atro  Switch  f^\ 


Right  Distributor. 
HIT  Starting 
..'  Magneto, 


FIG.  259. — Typical  wiring  system  for  12-cylinder  Vee  engine. 

system.  A  magneto  system  is  easier  for  the  pilot  to  operate; 
there  is  only  one  switch  handle  to  control  and  no  ammeter  to  be 
watched.  The  battery  system  has  distinct  advantage  when 
current  is  required  for  an  electric  starter,  lights  and  other  uses. 
The  firing  order  adopted  in  actual  engines  is  given  below. 


8-cylinder  90-deg.  Vee 


1L,  1R,  2L,  2R,  4L,  4R,  3L,  3R  (Curtis  OX5,  VX; 

Sunbeam  Arab) 

1L,  4R,  2L,  3R,  4L,  1R,  3L,  2R  (Hispano-Suiza) 
(counting  from  propeller) 


12-cylinder  60-deg.  Vee 


326  THE  AIRPLANE  ENGINE 

12-cylinder  45-deg.  Vee:  1L,  6R,  5L,  2R,  3L,  4R,  6L,  1R,  2L,  5R,  4L,  3R 

(Liberty) (counting  toward  propeller). 
1L,  2R,  5L.  4R,  3L,  1R,  6L,5R,  2L,  3R,  4L,  6R 
(Rolls-Royce  Falcon  and  Eagle) 
1L,  1R,  5L,  5R,  3L,  3R,  6L,  6R,  2L,  2R,  4L,  4R 
(Sunbeam  Maori,  Cossack) 
9-cylinder  Rotary:  1,  3,  5,  7,  9,  2,  4,  6,  8  (Gnome,  LeRhone,  BRl,  BR2, 

Clerget) 

6-cylinder  Vertical:  1,  5,  3,  6,  2,  4(Beardmore,  Galloway,  Siddeley,  Austro- 

Daimler,  Mercedes,  Fiat) 

7-cylinder  Radial:  1,  3,  5,  7,  2,  4,  6  (A.  B.  C.  Wasp) 
9-cylinder  Radial:  1,  3,  5,  7,  9,  2,  4,  6,  8  (A.  B.  C.  Dragonfly) 


CHAPTER  XIII 
LUBRICATION 

All  rubbing  surfaces  in  an  engine  should  be  lubricated.  The 
most  important  of  these  surfaces  are  the  cylinder  walls,  main 
bearings,  crankpins,  piston  pins,  and  camshaft  bearings,  but 
there  are  also  numerous  other  parts  to  be  lubricated. 

The  coefficient  of  friction  of  a  bearing  with  good  lubrication, 
moderate  pressures  and  high  speeds  is  practically  independent  of 
the  materials  composing  the  rubbing  surfaces,  but  is  proportional 
to  the  viscosity  of  the  oil,  to  the  rubbing  speed  and  to  the  area; 
it  is  independent  of  the  pressure.  For  high  pressures  and  low 
speeds  these  laws  do  not  hold;  for  velocities  from  100  to  500  ft. 
per  minute  the  coefficient  decreases  about  as  the  square  root  of 
the  velocity,  for  velocities  from  500  to  1,600  ft.  per  minute  it 
decreases  about  as  the  fifth  root  of  the  velocities,  while  above 
1,600  ft.  per  minute  it  is  practically  constant.  With  high  pres- 
sures the  coefficient  of  friction  increases. 

In  an  airplane  engine  the  loads  on  the  principal  bearing  surfaces 
are  variable,  going  through  a  cycle  of  changes  every  two  revolu- 
tions of  the  engine,  and  varying  from  a  maximum  to  a  low  mini- 
mum. For  example,  in  the  Liberty  engine  the  total  force  on  the 
crankpin  varies  from  4,980  to  1,500  lb.;  on  the  intermediate  main 
bearing  from  7,250  lb.  to  800  lb. ;  on  the  end  main  bearing  from  - 
4,025  lb.  to  zero;  on  the  center  main  bearing  from  7,700  lb.  to 
2,500  lb.;  the  piston  side  thrust  from  930  lb.  to  zero.  Further- 
more, the  direction  of  the  force  changes  in  these  principal  bearing 
surfaces,  so  that  the  portion  of  the  bearing  which  at  one  instant 
is  supporting  maximum  pressure  is  later  relieved  of  all  pressure. 
This  intermittent  application  of  the  load  is  favorable  to  good 
lubrication  and  permits  the  use  of  maximum  pressures  greatly 
in  excess  of-  what  would  be  possible  with  continuous  loading. 
The  oil  film  which  is  squeezed  out  by  the  application  of  the  maxi- 
mum pressure  is  replaced  during  the  reduction  or  reversal  of  the 
pressure. 

The  maximum  load  per  square  inch  of  projected  area  is  great- 
est on  the  piston  pin,  which  has  a  diameter  considerably  less  than 

327 


328  THE  AIRPLANE  ENGINE 

the  crankpin;  in  the  Liberty  engine  this  is  2,580  Ib.  per  square 
inch  as  against  932  Ib.  on  the  crankpin;  1,675  Ib.  on  the  center 
main  bearings;  1,580  Ib.  on  the  intermediate  main  bearings; 
815  Ib.  per  square  inch  on  the  end  main  bearings. 

The  rubbing  speeds  of  the  main  bearings  of  airplane  engines 
range  usually  from  16  to  20  ft.  per  second;  the  rubbing  speeds  at 
the  crankpins  will  be  somewhat  lower. 

The  total  friction  work  at  any  bearing  is  proportional  to  the 
product  of  the  mean  total  load  by  the  rubbing  speed.  The  limit- 
ing factors  for  a  bearing  for  continuous  operation  are  the  mean 
pressure  per  square  inch  of  projected  area  and  the  rubbing  speed. 
The  product  of  these  two  is  a  good  index  of  the  service  of  the 
bearing.  In  the  Liberty  engine  this  load  index  is  13,500  Ib.-ft. 
sec.  for  the  crankpin;  24,670  for  the  center  main  bearing;  13,650 
for  the  intermediate  main  bearing;  and  11,900  for  the  end  main 
bearings. 

The  permissible  pressure  on  the  bearing  depends  on  the  viscos- 
ity and  therefore  on  the  temperature  of  the  oil.  The  temperature 
tends  to  rise  and  must  be  kept  down  by  oil  cooling.  Tempera- 
tures of  160°F.  and  higher  are  common. 

The  lubrication  of  the  cylinder  offers  problems  quite  different 
from  the  lubrication  of  the  rest  of  the  engine.  The  side  thrust 
pressures  are  moderate;  the  piston  speed  is  high,  reaching  2,000  ft. 
per  minute,  or  33  ft.  per  second,  and  the  maximum  speed  (at 
mid-stroke)  is  about  52  ft.  per  second.  The  friction  work  under 
these  conditions  would  not  be  a  serious  charge  against  the  engine 
if  the  oil  film  could  be  maintained  in  good  condition,  but  as  pointed 
-out  on  page  24  the  viscosity  of  the  oil  film  on  the  cylinder  walls 
is  greatly  raised  by  carbonization.  The  oil  film  on  the  walls 
and  around  the  piston  rings  has  to  serve  another  purpose  besides 
acting  as  a  lubricant;  it  acts  as  a  seal  to  prevent  the  blowing  of  the 
gases  past  the  piston.  For  this  purpose  high  viscosity  is  useful. 

In  starting  cold,  in  idling  with  overrich  mixtures,  and  in  cold 
weather,  a  certain  amount  of  liquid  fuel  will  meet  the  cylinder 
walls  and,  being  perfectly  miscible  with  the  mineral  lubricating 
oil,  it  will  dilute  the  oil  film  and  the  thinned  oil  will  then  run  down 
the  cylinder  walls  and  dilute  the  oil  in  the  crankcase.  This 
phenomenon,  which  is  very  common  in  automobile  practice,  is 
not  so  usual  in  airplane  engines  because  of  the  higher  volatility 
of  the  aviation  fuels.  The  ordinary  gasoline  of  commerce  has  a 
high  end  point,  which  means  that  it  has  kerosene  constituents 


LUBRICATION  329 

which  will  not  vaporize  during  admission  and  will  be  deposited 
in  the  liquid  form  on  the  cylinder  walls.  With  aviation  fuel  the 
dilution  when  it  occurs  is  by  more  volatile  elements  which  tend 
to  vaporize  out  from  the  hot  body  of  oil  so  that  the  dilution  of 
the  crankcase  oil  is  not  cumulative  as  with  automobile  engines. 

The  friction  at  the  bearings  of  an  engine  results  in  heat  which 
has  to  be  taken  away  as  fast  as  it  is  generated  if  the  bearings  are 
not  to  rise  in  temperature.  Much  of  this  heat  is  conducted 
through  the  metal  to  cooler  parts  of  the  engine  but  part  is  carried 
away  by  the  oil  itself.  For  this  purpose  a  large  flow  of  oil  is 
desirable.  The  amount  of  oil  circulated  in  the  Liberty  engine  is 
about  12  gal.  per  minute  and  the  temperature  rise  may  average 
about  10°F.  This  corresponds  to  a  heat  abstraction  of  about 
45  B.t.u.  per  minute.  If  the  bearing  friction  work  is  taken  as 
1%  Ib.  per  square  inch  of  piston  area  (see  p.  24),  the  correspond- 
ing heat  generated  will  be  about  200  B.t.u.  per  minute.  The 
volume  of  oil  circulated  in  this  case  is  not  sufficient  for  complete 
cooling  of  the  bearings. 

Viscosity. — The  oil  which  gives  minimum  temperature  rise  of 
the  bearing  is  the  best  to  use,  other  factors  being  equal.  An  oil 
of  lower  viscosity  will  cause  greater  friction  because  it  will  squeeze 
out  and  allow  a  closer  contact  of  metallic  parts  and  increase 
metallic  friction  and  wear.  An  oil  of  higher  viscosity  results  in 
increased  fluid  friction  of  the  oil.  With  a  complete  film  of  oil, 
the  oil  flows  like  a  pack  of  playing  cards  sliding  over  each  other, 
the  outer  layers  adhering  to  the  surfaces  and  not  sliding  with 
reference  to  them.  The  actual  fluid  friction  Ft  in  pounds,  is  given 
by  the  equation 

PXAXV 
5,760  X  t 

where  P  is  the  absolute  viscosity  in  poises;  A  is  the  rubbing  area 
in  square  inches;  V  is  the  rubbing  velocity  in  feet  per  second; 
and  t  is  the  oil  film  thickness  in  inches.  This  formula  indi- 
cates that  the  friction  diminishes  as  the  thickness  of  the  film 
increases. 

The  measurement  of  viscosity  has  been  standardized  in  this 
country  and  is  determined  by  the  Saybolt  Universal  Viscosimeter 
in  which  the  fluid  to  be  analyzed  flows  through  a  tube  0.1765  cm. 
diameter,  1.225  cm.  long,  under  an  average  head  of  7.36  cm.,  from 
a  vessel  2.975  cm.  diameter.  The  time  in  seconds  required  for 
60  c.c.  of  oil  to  flow  through  the  tube  is  the  viscosity  in  seconds 


330  THE  AIRPLANE  ENGINE 

Saybolt.     The  absolute  viscosity  is  obtained  from  the  equation 
100  P  = 


where  G  is  the  specific  gravity  of  the  oil  and  S  is  the  Saybolt 
viscosity. 

One  of  the  most  troublesome  features  of  lubrication  is  the 
decrease  in  viscosity  of  oils  with  rise  in  temperature.  It  is  found 
that  if  the  logarithm  of  the  Saybolt  viscosity  is  plotted  against 
the  temperature  (Farenheit),  on  cross-section  paper,  the  points 
lie  close  to  a  straight  line  —  this  is  a  purely  accidental  relation. 

The  viscosities  of  the  principal  American  lubricating  oils  at 
temperatures  of  100°,  150°  and  212°F.  are  given  in  Table  16. 
It  will  be  noticed  that  the  very  heavy  oils  fall  off  most  rapidly  in 
viscosity  as  the  temperature  is  raised  so  that  whereas  the  range 
in  Saybolt  viscosities  of  the  oils  tabulated  at  100°F.  is  about  11  to 
1,  at  212°  it  is  only  3  to  1. 

The  desired  physical  characteristics  of  a  lubricating  oil  are  as 
follows  : 

(a)  Body  sufficient  to  prevent  metallic  contact  under  maximum  pressure 
and  maximum  temperature. 

(6)  Lowest  viscosity  in  keeping  with  the  above  conditions. 

(c)  Capacity  of  resisting  high  temperatures  without  decomposition. 

(d)  Fluidity  at  minimum  temperatures. 

(e)  High  fire  test. 

(/)  Freedom  from  oxidation. 

(g]  Freedom  from  corrosive  action  on  metals. 

The  standard  specifications  for  lubricating  oils  for  airplane 
engines  adopted  by  the  U.  S.  Army  and  Navy  are  given  below. 
Grade  1  is  the  Navy  specification,  Grade  2  the  Army.  The 
specifications  are  also  different  for  summer  and  winter  use. 

Flash  Point.  —  The  flash  point  of  Grade  1  shall  not  be  lower 
than  400°F.;  for  Grade  2  not  lower  than  500°F. 

Viscosity  at  210°F.  shall  be  within  the  following  limits: 

Grade  1  (summer)  ...........................     90-100  sec. 

Grade  1  (winter)  ............................     78-  85  sec. 

Grade  2  ....................................   125-135  sec. 

Pour  Test.—  Grade  1  not  above  45°F.  for  summer,  or  15°F. 
for  winter. 

Cold  Test.—  Grade  2  not  above  35°F. 


LUBRICATION 


331 


TABLE  16. — PROPERTIES  OF  REPRESENTATIVE  AMERICAN  LUBRICATING  OILS 
FOR  USE  IN  INTERNAL  COMBUSTION  ENGINES 


Kind  of  oil 

Physical  properties 

Baume 
grav. 

Flash,  deg.  F. 

Deg.F. 
burn 

Deg.  F. 
chill 

Viscosity,  seconds 

Open 
cup 

Closed 
cup 

100° 
F. 

150° 
F. 

212° 
F. 

Havoline: 
Light                            .... 

25.9 
25.0 
25.6 

28.1 
21.8 
26:3 
23.3 
21.1 
25.8 

27.6 
26.0 
28.9 

24.7 

29.1 
24.9 
29.3 

24.3 
25.4 

21.3 
20.9 
19.3 

26.2 
27.1 
26.3 
27.6 
24.7 
24.7 
26.2 
26.1 

28.6 
27.6 
15.0 

370 
385 
395 

370 
360 
500 
370 
370 
460 

360 
375 
430 
465 

400 
390 
420 

395 
385 

335 
350 
356 

455 
450 
435 
440 
440 
460 
410 
465 

415 
485 

380 
395 
410 

380 
360 
470 
380 
585 

360 
370 

445 
425 

410 
400 
430 

410 
380 

340 
350 
360 

450 
445 
435 
430 
450 
460 

460 

... 

430 
450 
455 

420 
420 
580 
425 
430 
540 

410 
430 
505 
535 

470 
450 
495 

470 
450 

380 
400 
420 

535 
530 
520 
515 
520 
540 
480 
550 

475 
550 

33 
34 
46 

0 
24 
41 
6 
8 
46 

20 
23 
34 

58 

26 
32 
40 

SI  at  0 
35 

SI  at  0 
SI  at  0 
10 

39 
38 
38 
34 
28 
27 
33 
44 

32 
46 

173 
237 
361 

167 
330 
1,640 
221 
300 
926 

140 
289 
340 
1,583 

181 
243 
316 

219 
300 

205 
301 
495 

795 
814 
517 
513 
413 
474 
329 

334 
1,196 
1,270 

66 
80 
111 

66 
97 
397 
74 
87 
243 

60 
95 
108 
356 

71 
81 
103 

77 
103 

69 
85 
119 

212 
222 
149 
151 
135 
134 
107 
355 

108 
300 
305 

42 
46 
54 

44 
49 
122 
45 
46 
86 

41 
50 
55 
110 

45 
47 
54 

46 
51 

42 
46 
51 

78 
80 
63 
64 
55 
58 
52 
111 

52 
100 
90 

Mobiloil: 
"E"  Light              

"  A  "  Medium         

"B"  Heavy           

Arctic  Lt    Med  

Arctic  Medium  
"BB"  Med.  Heavy  

Monogram: 
Light  
Medium   

Heavy  
Ex.  Heavy  

Perfection: 
"A"  Light 

"C"  Heavy 

Socony: 
Zero 

Texaco: 
Light                        

Medium                    

Heavy                   

Veedol: 
Aero  No.  1  

Aero  No.  2  

Aero  No.  3  

Aero  No   6 

Zero  Heavy  
Zero  Extra  Heavy  

Wolf's  Head: 
Heavy  

No   8 

332  THE  AIRPLANE  ENGINE 

Acidity. — Not  more  than  0.10  mg.  of  potassium  hydroxide 
shall  be  required  to  neutralize  1  gram  of  Grade  1  oil. 

Emulsifying  Properties. — The  oil  shall  separate  completely 
in  1  hr.  from  an  emulsion  with  distilled  water  at  a  temperature  of 
180°F. 

Carbon  Residue  in  Grade  1  shall  not  be  over  1.5  per  cent; 
in  Grade  2  not  over  2.0  per  cent. 

Precipitation  Test. — When  5  c.c.  of  the  oil  is  mixed  with 
95  c.c.  of  petroleum  ether  and  allowed  to  stand  for  24  hr.  it 
shall  not  show  a  precipitate  or  sediment  of  more  than  0.25  c.c. 

The  oil  shall  not  contain  moisture,  sulphonates,  soap,  resin 
or  tarry  constituents. 

The  Flash  Test  shows  the  temperature  at  which  the  vapor  from  a  sample, 
heated  in  an  open  cup,  will  ignite.  It  has  some  relation  to  loss  by  evapora- 
tion. The  open  cup  is  2^  in.  diameter,  1%6  in.  high  and  is  filled  to  within 
%  in.  of  the  top.  It  is  placed  on  a  metal  plate  and  heated  so  that  its  tem- 
perature rises  not  less  than  9°  nor  more  than  11°F.  per  minute.  A  test 
flame  ^2  in.  in  diameter  is  passed  across  the  top  of  the  cup,  taking  1  sec.  for 
the  passage,  at  every  5°F.  rise  of  temperature  of  the  oil.  The  flash  point  is 
the  temperature  at  which  a  flash  appears  at  any  point  on  the  surface  of  the 
oil.  Drafts  must  be  avoided. 

The  Viscosity  Test  has  been  discussed  on  page  329. 

The  Pour  Test  indicates  the  temperature  at  which  a  sample  of  the  oil 
will  just  flow.  The  oil  is  placed  in  a  glass  jar  \Y±  in.  diameter  and  4  to  5  in. 
long  to  a  depth  of  about  Y±  in.  and  the  jar  is  corked.  It  is  then  placed  in  a 
freezing  solution  and  at  each  5°F.  drop  in  temperature  it  is  taken  out  and 
tilted.  The  pour  test  is  5°  higher  than  the  temperature  at  which  the  oil  will 
not  flow  when  the  jar  is  placed  in  the  horizontal  position. 

The  Cold  Test  has  a  similar  purpose  to  the  Pour  Test  but  in  this  case  the 
oil  is  first  frozen  and  is  then  stirred  with  a  thermometer  until  it  will  run  from 
one  end  of  an  ordinary  4-oz.  sample  bottle  to  the  other.  The  temperature 
reading  at  that  time  is  the  Cold  Test. 

Carbon  residue  is  obtained  by  heating  10  grams  of  oil  in  a  porcelain  crucible 
placed  inside  two  iron  crucibles  with  covers.  It  is  heated  so  as  to  maintain 
a  vapor  flame  of  specified  length  and  heated  further  after  the  vapors  cease 
to  come  off.  After  cooling  the  weight  of  carbon  residue  in  the  porcelain 
crucible  is  determined. 

Reclaiming  Oil. — The  oil  used  in  a  stationary  airplane  engine 
not  only  becomes  diluted  by  the  heavier  constituents  of  the  fuel 
but  also  becomes  dirty  by  the  accumulation  of  free  carbon  from 
the  cylinder  walls,  of  metal  particles  worn  off  the  bearing  surfaces, 
and  of  other  solid  impurities.  The  oil  usually  has  to  be  changed 
between  the  fifth  and  twentieth  hour  of  flying  service;  most  oil 
is  not  used  more  than  5  hr.  It  can  be  reclaimed  by  allowing  it  to 


LUBRICATION  333 

stand  30  hr.  in  a  tall  bucket,  decanting  off  the  upper  two-thirds, 
filtering  and  warming  to  150°F.  A  more  thorough  process  is  to 
put  the  oil  with  some  water  and  soda  ash  in  a  steam-jacketed 
tank,  raising  the  temperature  to  212°F.,  forming  an  emulsion 
and  obtaining  a  precipitation  of  carbon,  iron,  and  dirt  after  a 
period  of  rest.  The  steam  drives  off  the  2  or  3  per  cent  of  diluent 
coming  from  the  volatile  aviation  gasoline.  A  recovery  of  85 
per  cent  is  possible  in  this  way  and  the  reclaimed  oil  is  at  least 
as  good  as  new  oil. 

A  centrifugal  oil  cleaner  has  been  tried  on  the  Liberty  engine 
with  considerable  success.  This  consists  of  a  spun  copper  bowl, 
5  in.  diameter,  rotating  at  IJ-^  crankshaft  speed;  the  centrifugal 
force  at  2,550  r.p.m.  is  about  45  times  that  of  gravity.  The  oil  is 
led  into  the  center  of  the  bowl  and  is  thrown  out  at  the  top. 
Examination  of  the  contents  of  the  bowl  after  a  run  show  that  it 
collects  metal  particles,  sand,  carbon,  rubber  and  other  solids. 
Its  use  should  increase  considerably  the  periods  between  changes 
of  oil  and  should  prolong  the  life  of  all  bearing  surfaces  by 
preventing  their  abrasion  by  solid  particles  in  the  oil. 

Castor  oil  is  employed  in  rotary  engines  in  which  the  gasoline  is 
admitted  to  the  crankcase  on  its  way  to  the  cylinders.  Mineral 
oil  (petroleum)  cannot  be  used  in  this  case  as  it  is  miscible  with 
gasoline  and  its  use  would  result  in  a  thinning  of  the  lubricant 
and  a  wastage  of  fuel. 

Methods  of  Lubrication. — The  splash  system  of  lubrication 
often  employed  in  automobile  engines  is  not  satisfactory  for 
airplane  engines  on  account  of  the  high  loading  at  which  they  are 
operated  and  also  because  of  the  extreme  variations  in  engine 
orientation  during  flight.  For  the  last  reason  also  a  wet  sump  is 
undesirable  since  it  will  deluge  some  of  the  cylinders  during  such 
airplane  evolutions  as  a  nose  dive  and  may  result  in  trouble  from 
excess  of  oil  in  the  cylinder.  Consequently  the  modern  airplane 
engine  is  provided  with  a  pressure  oiling  system  and  a  sump, 
which  is  usually  kept  dry  by  a  scavenger  pump. 

The  normal  lubricating  system  is  as  follows.  The  scavenger 
pump  or  pumps  take  oil  from  the  sump  and  deliver  it  to  the 
external  oil  tank.  The  pressure  pump  takes  oil  from  the  oil 
tank  and  discharges  it  into  a  distributing  main  from  which 
branches  go  to  each  of  the  main  bearings.  Oil  enters  some  of 
the  hollow  main  journals  through  small  holes  which  register 
with  corresponding  holes  or  channels  in  the  bearing  and  there- 


334  THE  AIRPLANE  ENGINE 

by  with  the  branch  oil  pipes.  Usually,  alternate  journals  are 
rilled  with  oil  and  the  crank  cheeks  on  the  two  sides  of  these 
journals  are  drilled  to  connect  with  the  hollow  crankpins  and 
thereby  permit  lubrication  of  the  crankpins  through  appropriate 
holes  in  them.  The  oil-containing  journals  and  all  the  crankpins 
have  closed  ends  (see  p.  144).  The  lubrication  of  the  piston  pin  is 
sometimes  carried  out  by  oil  pipes  running  along  the  connecting 
rod  and  registering  every  revolution  with  the  oil  hole  in  the  crank- 
pin;  in  other  cases  the  oil  thrown  out  by  centrifugal  force  from  the 
crankpin  is  relied  on  to  lubricate  the  piston  pin  as  well  as  the 
cylinder  wall. 

The  camshaft  is  lubricated  from  an  oil  pipe  from  the  end 
of  the  distributing  main,  connecting  with  it  usually  by  an  annular 
groove  in  the  front  main  bearing.  The  camshaft  is  hollow  and 
acts  as  an  oil  carrier  discharging  oil  through  small  holes  at  each 
bearing.  The  oil  escaping  from  the  bearings  lubricates  the  cams 
and  returns  to  the  crankcase  over  the  distributing  gears,  meeting 
there  the  oil  escaping  from  the  main  bearings,  crankpins  and 
cylinders. 

The  lubricating  system  of  the  Liberty  engine  follows  the  lines 
indicated  above.  The  cylinders,  pistons  and  wristpins  are 
lubricated  by  oil  spray  from  the  crankpin.  A  double  scavenging 
pump  at  the  rear  of  the  engine  (Fig.  261)  keeps  the  two  sumps 
drained  and  returns  the  oil  to  the  outside  oil  tank.  In  the 
Hispano-Suiza  engine  (Fig.  51)  a  dry  sump  is  also  used.  A 
single  gear-type  scavenging  pump  is  driven  directly  from  the 
rear  of  the  crankshaft;  an  eccentric  sliding-vane  pressure  pump 
is  mounted  on  the  vertical  water- pump  shaft  and  rotates  at  1.2 
times  the  crankshaft  speed.  The  vane  pump  forces  the  oil 
through  a  filter  to  the  main  oil  pipe  in  the  lower  crankcase. 
There  are  four  oil  holes  in  each  crank  pin  through  which  oil  goes 
to  the  bearing  and  is  thrown  off  to  lubricate  the  cylinder  and 
wrist  pin.  A  small  hole  is  provided  in  the  leading  face  of  each 
cam  to  lubricate  the  cam  and  its  follower. 

In  the  Curtiss  K-12  (Fig.  55)  the  pumps  are  located  in  the 
lower  part  of  the  crankcase  and  are  driven  through  a  horizontal 
shaft.  There  is  a  triple-gear  scavenging  pump  and  two  pressure 
pumps  arranged  in  a  unit  with  the  spiral  driving  gears  and 
surrounded  by  a  filtering  screen.  The  oil  is  supplied  to  the  main 
bearings  in  the  usual  way  and  is  then  conveyed  to  the  crankpins 
through  small  tubes  built  into  the  crankshaft.  The  oil  is  cooled 


LUBRICATION  335 

by  a  temperature  regulator  through  which  the  jacket  water 
circulates  on  its  way  from  the  radiator  to  the  pump. 

The  Curtiss  OX  engine  uses  a  wet  sump  (Fig.  59)  covered 
by  an  oil-pan  partition.  A  single-gear  pump  driven  from  a  beveled 
gear  on  the  crankshaft  at  the  propeller  end  sucks  oil  from  the 
sump  and  discharges  it  into  the  rear  end  of  the  hollow  camshaft 
whence  it  goes  through  tubes  to  the  crankshaft  bearings.  In  this 
engine  there  is  a  continuous  closed  oil  passage  from  one  end  of  the 
crankshaft  to  the  other.  The  cylinder  is  lubricated  by  oil  spray. 
The  oil-pan  partition  has  a  half-inch  hole  at  its  center  for  the 
return  of  oil  to  the  sump. 

The  Hall-Scott  L-6  engine  (Fig.  63)  uses  a  wet  sump  and  has 
scavenger  and  pressure  pumps  mounted  as  a  unit  inside  the  lower 
crankcase  and  driven  through  an  inclined  shaft  from  a  bevel  gear 
which  is  at  the  rear  of  the  crankshaft.  The  system  is  otherwise 
similar  to  the  Liberty  engine.  An  oil  sight  gage  (Fig.  64) 
shows  the  level  in  the  sump.  Splash  plates  in  the  lower  case  pre- 
vent excessive  splash  from  the  dipper  action  of  the  connecting 
rods. 

In  the  Napier  "Lion"  engine  three  oil  pumps  are  combined  as 
a  single  unit  at  the  extreme  rear  of  the  engine ;  they  are  of  the  gear 
type  and  are  driven  at  half  engine  speed.  The  suction  pumps  draw 
the  oil  away  from  the  two  ends  of  the  lower  crankcase  by  two 
separate  steel  pipes  (Fig.  72)  and  discharge  to  the  tank  through 
a  common  pipe.  The  pressure  pump  delivers  to  both  ends  of  the 
crankshaft  and  to  the  three  camshaft  casings.  There  is  a  con- 
tinuous passage  for  oil  through  the  crankshaft. 

In  the  Fiat -650  engine  (Fig.  76)there  are  scavenger  pumps  at 
the  two  ends  of  the  lower  crankcase  and  a  pressure  pump  di- 
rectly under  the  rear  scavenger  pump.  They  are  driven  by  bevel 
gears  from  the  horizontal  tubular  shaft  inside  the  crank  casing 
and  are  mounted  in  ball  bearings  as  well  as  the  horizontal  shaft 
and  the  spur  gear  which  drives  it.  The  oil  drawn  from  the  main 
tank  is  discharged  into  a  copper  main  cast  into  the  crankcase. 
The  main  bearings  are  fed  from  copper  branch  pipes.  In  other 
respects  the  system  is  normal. 

In  the  Benz-230  (Fig.  77)  a  wet  sump  is  used  and  the  triple 
oil  pump  is  submerged  in  the  sump.  The  main  pressure  pump  A 
(Fig.  260)  draws  oil  from  the  reservoir  in  the  sump  and  discharges 
it  to  the  main  bearings  through  a  distributing  main  and  branch 
pipes.  The  supply  to  the  piston  pins  is  through  small  pipes 


336 


THE  AIRPLANE  ENGINE 


inside  the  tubular  connecting  rods.  Fresh  oil  is  fed  into  the 
sump  by  the  small  suction  pump,  B,  from  the  oil  tank  while 
the  correct  working  oil  level  is  maintained  in  the  reservoir  by  the 
pump,  C,  whose  curved  suction  pipe  (see  Fig.  77)  terminates  at 
the  desired  oil  level.  All  return  oil  passes  over  the  transverse 
air  pipes  in  the  lower  crankcase  and  is  cooled  by  them. 

In  the  Maybach  engine  (Fig.  80)  scavenger  pumps  are 
mounted  at  both  ends  of  the  crankcase  and  the  pressure  pump  is 
placed  behind  the  rear  scavenger  pump.  All  three  are  operated 
by  the  same  horizontal  shaft  driven  by  spur  gearing  from  the 
front  end  of  the  engine  shaft.  The  oil  is  discharged  into  an 


Return  Pipe  to  Main  Dtlivtry  Pips 

Tank  from         /^\  fromTankto  PumpB 
PumpC-... 


Oil  L  eve  fin  Sump. 

K. 


JoTank 


From  Tank       From  Sump  ToTank   / 
fB  I  Q.   ;        r?,. 


FromSu. 


Delivery  to 
Main  Bearings 


From  Sump 
to  PumpC 


To  Sump 


A- Pump  Supply  ing 
Main  Bearin9 
From  Sump 


B-  Pump  Supplying  C-Pump 
Sump  from  Returning 

Oil  Tank  Oil  from  Sump 

To  Tank 


From  Sump      Supply  to 
fv  Pump-A       Sump  from 

Pump-B 
D-  End  View 
of  Pump 
Cover. 


FIG.  260. — Triple-gear  pump  of  Benz  engine. 

external  oil  main  on  the  upper  crankcase  and  past  individual 
screens  into  branch  pipes  drilled  through  the  transverse  webs 
into  the  main  bearings.  The  oil  thrown  off  from  these  bearings 
is  caught  in  aluminum  scoops  bolted  to  both  ends  of  each  crank- 
pin,  is  carried  by  centrifugal  force  into  the  hollow  crankpin  and 
thence  through  radially  bored  holes  to  the  bearing.  The  piston 
pins  are  oiled  through  the  internal  pipes  in  the  connecting  rods. 
Baffle  plates  bolted  to  the  upper  crankcase  just  below  the  cylin- 
ders prevent  an  excess  of  oil  reaching  the  cylinders. 

Relief  Valves. — All  pressure-feed  systems  are  provided  with  a 
relief  valve  on  the  discharge  side  of  the  pump.  This  is  a  spring- 
loaded  valve  as  in  Fig.  261  and  is  set  for  the  maximum  allowable 
pressure.  The  oil  pressure  is  high  in  starting  especially  in  cold 


LUBRICATION  337 

weather  when  the  viscosity  of  the  oil  is  very  great.  The  normal 
operating  oil  pressures  after  fully  warming  up  are  about  25  to  30 
Ib.  per  square  inch  in  Liberty  engines,  50  to  60  Ib.  in  Curtiss 
engines,  40  to  65  Ib.  in  Hispano-Suiza  engines.  In  cold  weather 
it  is  best  to  drain  off  the  oil  after  a  flight  and  to  fill  up  with 
hot  oil  before  starting. 

The  location  of  oil  grooves  in  the  bearings  is  a  matter  of  con- 
siderable importance  on  which  there  is  much  divergence  of 
practice.  The  actual  pressure  on  the  oil  film  will  vary  from  zero 
at  the  ends  of  the  bearings  and  at  the  split  of  the  bearing  to 
possibly  as  much  as  10,000  Ib.  per  square  inch  in  the  center  of  the 
loaded  area  at  the  moment  of  maximum  loading.  As  the  oil 
pressure  does  not  exceed  50  Ib.  per  square  inch  it  is  obvious  that 
oil  cannot  be  forced  in  at  the  place  of  maximum  loading  unless  the 
pressure  at  that  place  falls  below  50  Ib.  per  square  inch  during 
some  part  of  the  cycle.  There  should  be  no  oil  grooving  length- 
wise in  the  middle  of  the  most  loaded  half  of  a  bearing;  such 
grooves  are  channels  of  escape  for  the  oil  and  may  result  in  such 
thinning  of  the  film  as  to  increase  the  friction  and,  possibly,  to 
cause  seizing.  The  most  heavily  loaded  part  of  the  main  bearing 
is  usually  the  middle  of  the  lower  cap  and  it  is  at  this  place  that 
the  oil  usually  enters.  The  grooves  should  then  be  two  helical 
grooves  intersecting  at  the  oil  hole  at  the  center  of  the  bottom 
half  of  the  bearing  and  running  to  the  split  but  not  too  near  the 
ends  of  the  bearing.  Short  helical  grooves  in  the  upper  half  of 
the  bearing  may  start  at  the  split  opposite  the  lower  grooves  but 
should  not  go  more  than  half-way  up  each  side.  Similar  groov- 
ing should  be  provided  at  the  crankpin  bearing,  which  also  is  most 
heavily  loaded  at  the  lower  half.  .  - 

Wherever  practicable  it  is  desirable  that  the  oil  should  enter  at 
the  place  of  minimum  average  bearing  pressure.  It  is  the 
cyclical  variation  in  the  loading  that  makes  possible  the  proper 
lubrication  of  the  heavily  loaded  bearing  surfaces  of  aviation 
engines.  The  oil  pressures  employed  are  in  themselves  not  nearly 
adequate  to  support  the  loads  but  are  'required  to  overcome  the 
viscous  and  frictional  resistance  to  the  flow  of  the  oil  to  the  va- 
rious bearings  and  also  to  ensure  that  the  oil  channels  will  clear 
themselves  of  small  obstructions. 

The  amount  of  oil  circulated  per  minute  is  determined  not 
only  by  the  lubrication  needs  but  also,  as  pointed  out  on  page  329, 
by  the  extent  to  which  the  oil  is  used  as  a  cooling  medium.  For 

22 


338  THE  AIRPLANE  ENGINE 

lubrication  the  amount  should  probably  be  some  function  of  the 
total  projected  areas  of  the  main  bearings  and  crankpins; 
expressed  in  this  way  the  use  of  oil  varies  from  0.1  to  0.5  Ib.  per 
square  inch  of  projected  bearing  area  per  minute.  In  terms  of 
the  power  delivered  by  the  engine  the  oil  circulated  varies  from 
0.025  to  0.15  Ib.  per  horse-power  minute. 

The  oil  consumption  of  an  engine  as  usually  measured  is  the 
amount  of  oil  which  has  to  be  added  to  the  system  to  make  up  for 
oil  burned  or  otherwise  used  up  during  the  engine  operation. 
This  quantity  varies  from  0.02  to  0.05  Ib.  per  horse-power  hour 
in  stationary  water-cooled  engines  but  may  go  as  high  as  0.15 
Ib.  in  rotaries. 

Oil  Pumps. — The  great  majority  of  airplane  engines  use  gear 
pumps  both  for  scavenging  and  pressure  pumps.  A  simple  gear 
pump  consists  of  a  power-driven  spur  gear  meshing  closely  into  an 
exactly  similar  driven  gear,  the  gears  being  enclosed  in  a  casing 
with  the  minimum  working  clearance  above  and  below  the  gears 
and  also  around  them  except  where  they  mesh.  The  oil  inlet  is  on 
the  side  where  the  gears  separate;  the  discharge  is  on  the  side 
where  the  gears  meet.  If  there  were  no  leakage  the  oil  carried 
from  the  inlet  to  the  discharge  side  would  be  equal  to  the  space 
between  the  teeth,  but  some  of  this  is  brought  back  to  the  inlet 
side  since  a  tooth  going  into  mesh  does  not  fill  the  space  between 
adjacent  teeth  and  consequently  does  not  displace  all  the  oil 
content  of  that  space.  The  capacity  of  each  gear  can  be  taken 
approximately  as  equal  to  half  the  annular  space  between  the 
roots  and  tips  of  the  gear,  or,  for  the  two  gears,  as  equal  to  the 
whole  of  the  annular  space  of  one  gear.  As  the  width  of  this 
annular  space  increases  with  decrease  of  the  number  of  teeth 
(decrease  of  pitch)  it  is  evident  that,  for  a  given  pitch  diameter, 
capacity  can  be  increased  by  decreasing  the  pitch. 

Two  scavenger  gear  pumps  are  sometimes  combined  into  a 
triple-gear  pump  as  in  the  Liberty  engine.  In  this  case  the 
driving  gear  is  central  and  the  inlets  to  the  driven  gears  are  on 
opposite  sides  of  the  pump. 

The  driving  gears  for  scavenger  and  pressure  pumps  are  usually 
placed  close  to  one  another  and  driven  by  the  same  shaft.  In  the 
Benz  engine  (Fig.  260)  three  such  gear  pumps  are  mounted  on  the 
driving  shaft  and  function  as  indicated  in  the  figure.  The 
Liberty  oil  pump  is  driven  at  one  and  one-half  times  crankshaft 
speed  and  has  a  capacity  of  1.9  gal.  per  minute  at  normal  speed. 


LUBRICATION 


339 


The  pump  consists  of  a  double  scavenging  pump  with  three 
gears,  A,  B  and  C,  Fig.  261,  drawing  oil  from  the  two  sumps  and 


Plan  Vi'ew  of)  Oil  Pump 

Engine  Sump-Oilfo  Tank 

s~fs*L'  from  Engine 


Oil  from  Engine  Sump 
to  Upper  Pocket  - 

' 

Gears  in 
Lower  Pocket , 


If 


'n  Upper 
Pocket 

..  Oil  from  Lower 
/  Poctef  Under 
Pressure 


Oil  from  Tank 


Dram  Plug 
Cnoss-sec-hon  of  Liberf^-12  Oil  Pump 
FIG.  261. — Gear  pump  of  Liberty  engine. 

a  pressure  pump  immediately  below  it  giving  an  oil  pressure  of 
35  to  50  Ib.  per  square  inch  at  engine  speeds  from  1,500  to  1,800 
r.p.m.  The  pressure  gears,  A',  C',  (Fig.  261)  are  immediately 


340 


THE  AIRPLANE  ENGINE 


below  the  gears  A  and  C  of  the  scavenging  pump.  The  gear 
A'  is  driven  from  the  pump  shaft  and  the  upper  train  is  operated 
through  a  vertical-shaft  connection  from  Cf  to  C.  The  pressure 
pump  draws  oil  from  the  tank  through  a  copper  pipe,  the  oil 
passing  through  the  large  lower  strainer  and  entering  the  gear 
housing  at  M ;  it  is  discharged  through  the  passage  NOP  to  the 
distributing  manifold.  The  pressure  relief  valve  is  shown  in 


1800  2000 

Rev  per  Min  of  Engine 

FIG.  262. — Performance  curves  of  gear  pump  of  Packard-180. 

Fig.  261;  the  spring  is  usually  set  for  a  pressure  of  50  Ib.  per 
square  inch.  The  discharge  from  the  relief  valve  goes  to  the 
suction  side  of  the  pressure  pump. 

The  oil  from  the  rear  sump  entering  the  scavenging  pump  after 
passing  through  the  upper  strainer  is  drawn  through  E  and  dis- 
charged to  the  outlets  F  and  G,  which  are  connected  by  a  passage 
in  the  pump  body.  From  G}  the  oil  goes  through  the  passages  K 


FIG.  263. — Sliding-vane  pump. 

and  L  to  the  oil  tank.  Oil  from  the  front  sump  is  taken  through 
an  internal  pipe  and  the  passage  HIJ  and  is  also  discharged  at 
F  and  G. 

The  volumetric  efficiency  of  the  gear  pump  can  be  made  very 
high — probably  up  to  90  per  cent;  the  over-all  efficiency  is  50  to  60 
per  cent.  Tests  of  the  pressure  pump  of  the  Packard  180-h.p. 
engine  give  the  results  shown  in  Fig.  262.  It  will  be  seen  that 
the  slip  is  very  low  since  the  discharge  curves  are  almost  the 


LUBRICATION 


341 


From 
Sumo 


FIG.  264. — Plunger  pump  of  Basse-Selve  engine. 


Oil  Tank 


Main  Pressure  Pumps A 

Scavenger  Pumps B 

Delivery  of  Pumps  A  .. 

Suction  of  Pumps    A      -  —  —  — 
Defivery  of  Pumps    B 
Suction  of  Pumps     B 
Gravity  Feed  from  Oil  Tank 
Oil  Radiator 
Pulsation  Damper 


FIG.  265. — Oil  system  of  Basse-Selve  engine. 


342 


THE  AIRPLANE  ENGINE 


same  with  free  discharge  and  with  discharge  against  pressures 
which  vary  from  48  Ib.  at  1,200  r.p.m.  to  58  Ib.  at  2,000  r.p.m. 
It  may  be  further  noted  that  the  slopes  of  the  discharge  curves  up  to 
1,600  r.p.m.  are  such  as  to  go  through  the  zero  of  ordinates; 
that  is,  the  pump  discharge  is  directly  proportional  to  its  r.p.m. 
The  eccentric  sliding-vane  pump  (Fig.  263)  is  used  occasionally 
but  is  being  displaced  by  the  gear  type.  The  sliding  vanes  are 


To  De/ivery 
Main ••• 


To  Oil  Tank 
From  Sump 


FIG.  266. — Plunger  pumps  of  Salmson  engine. 

pressed  out  by  springs  in  a  slotted  cylinder  which  is  mounted 
eccentrically  in  a  cylindrical  casing  and  is  power-driven. 

Plunger  pumps  are  used  in  some  of  the  German  engines,  being 
operated  by  eccentrics  or  by  scroll  cams.  In  the  Basse-Selve 
engine,  (Fig.  264)  the  oil  pump  is  driven  by  a  worm  gear.  The 
pump  consists  of  two  double-acting  steel  plungers  which  work 
vertically  in  aluminum  barrels  and  make  in  effect  four  pumps. 


FIG.  267. — Action  of  LeRhone  oil  pump. 

The  plungers  are  rotated  by  the  worm  gear  and  are  simultan- 
eously reciprocated  by  the  action  of  the  scroll  cam  cut  in  the 
spindle  and  operated  by  a  hardened  steel  roller  working  on  a  pin 
screwed  into  the  pump  body.  At  each  stroke  of  the  two  plungers 
oil  is  drawn  from  the  cooler  tank  t6  one  of  the  inner  pump  cham- 
bers and  is  discharged  from  the  other  inner  pump  chamber  into 
the  delivery  main.  At  the  same  time  oil  is  sucked  out  of  one  of 
the  engine  sumps  into  one  of  the  outer  pump  chambers  and  is 


LUBRICATION 


343 


pumped  from  the  other  outer  pump  chamber  into  the  cooler 
tank.  A  diagrammatic  view  of  the  system  is  shown  in  Fig.  265. 
An  entirely  different  arrangement  of  plunger  pumps  is  used 
in  the  Salmson  engine  (Fig.  266).  In  this  case  two  plungers  are 
used,  the  larger  one  being  the  scavenging  pump.  The  plungers 
are  pivoted  on  a  crankpin  rotated  by  worm  gearing.  The  pump 
bodies  are  pivoted  and  oscillate  through  a  small  angle  as  the 
crankpin  rotates  and  the  plungers  make  their  strokes.  The 
connections  to  suction  and  discharge  are  made  when  openings 
in  the  oscillating  pump  bodies  register  with  appropriate  openings 


Side      View 
FIG.  268.— Mounting  of  oil  tank. 

in  the  pump  casing.  The  method  of  action  of  a  single  pump  of 
this  type,  used  on  the  Le Rhone  engine,  is  shown  diagrammatically 
in  Fig.  267.  The  port,  P,  in  the  oscillating  cylinder  comes 
alternately  opposite  the  intake  port,  /,  and  the  delivery  port,  D. 
The  method  of  mounting  the  oil  tank  and  of  connecting  it  to 
the  engine  in  the  USD-9A  airplane  is  shown  in  Fig.  268.  The 
tank  is  slung  from  the  engine  sills,  L,  and  is  located  just  below  the 
crank  case,  C.  The  tube,  A ,  is  the  oil  return  pipe  from  the  scaven- 
ger pump  to  the  tank;  the  supply  pipe  from  the  tank  to  the 
pressure  pump  is  shown  in  dotted  lines.  The  front  sump  con- 
nects by  a  pipe,  B,  to  a  Y-shaped  air-pressure-relief  pipe  inside 
the  tank,  the  upper  ends  of  which  extend  nearly  to  the  top  of  the 
tank.  This  acts  as  a  pressure-relief  vent  for  the  tank. 
The  oil  is  cooled  by  longitudinal  air  pipes,  H. 


CHAPTER  XIV 
THE  COOLING  SYSTEM 

All  airplane  engines  are  ultimately  air-cooled.  The  only 
option  is  as  to  whether  the  air  shall  be  applied  directly  or  indi- 
rectly. In  the  latter  case,  the  heat  is  removed  by  water  which 
is  then  cooled  by  the  air.  Indirect  (water)  cooling  offers  two 
advantages:  (1)  the  transmission  of  heat  from  the  cylinder  is 
more  rapid  to  water  than  to  air,  and  (2)  the  ultimate  cooling 
surface  (in  the  radiator)  can  be  made  much  greater  than  is 
possible  at  the  cylinder.  Direct-air  cooling  has  the  advantage  of 
reduced  total  weight  and  diminished  vulnerability;  a  bullet  hole 
through  the  radiator  or  jacket  will  put  the  whole  engine  out  of 
action  while  a  hole  through  one  cylinder  of  an  air-cooled  engine 
may  put  that  cylinder  alone  out  of  action. 

Air  cooling  in  airplanes  is  used  almost  exclusively  in  rotary 
and  radial  engines.  In  vertical  or  Vee  engines  with  several 
cylinders  in  line,  the  cooling  problem  is  much  more  difficult, 
although  it  has  been  met  by  the  use  of  suitable  cowling  to  direct 
the  air  on  to  the  different  cylinders.  With  the  rotary  and  radial 
types  the  motion  of  the  plane  and  the  location  of  the  engine  in  the 
slip  stream  of  the  propeller  ensure  an  adequate  flow  of  air  for 
cooling;  with  the  multi cylinder  vertical  or  Vee  type  it  is  some- 
times necessary  to  add  a  fan  to  improve  the  air  circulation. 

The  resistance  or  drag  of  air-cooled  cylinders  is  considerable 
but  has  not  been  determined  experimentally  in  a  satisfactory 
manner.  In  the  rotary  engine  there  is,  in  addition  to  the  drag, 
the  resistance  due  to  churning  which  reduces  directly  the  b.h.p. 
of  the  engine.  This  resistance  increases  so  rapidly  with  speed 
that  it  is  not  found  desirable  to  operate  rotary  engines  at  speeds 
in  excess  of  about  1,400  r.p.m.;  the  increase  in  indicated  power 
which  results  from  increased  speed  is  largely  used  up  in  over- 
coming the  increased  air  resistance.  The  total  work  done  by  an 
air-cooled  engine  in  overcoming  air  resistance  is  probably  greater 
under  ordinary  conditions  of  operation  than  the  total  work  done 
by  a  water-cooled  engine  in  overcoming  the  drag  of  its  radiator. 
Until  quite  recently,  air-cooled  cylinders  were  at  a  great  dis- 
advantage both  in  fuel  economy  and  in  power  developed  per 

344 


THE  COOLING  SYSTEM  345 

unit  volume  of  piston  displacement,  but  recent  constructions 
have  put  the  air-  and  water-cooled  engines  very  nearly  on  a  par 
in  these  respects.  They  have  always  had  an  advantage  in 
weight  per  horse  power. 

The  heat  which  has  to  be  removed  from  the  cylinder  in  order  to 
keep  the  temperature  of  the  cylinder  within  the  limit  which 
permits  satisfactory  operation  of  the  engine  is  usually  about 
equal  to  the  heat  equivalent  of  the  work  done  in  the  cylinder,  or 
42  B.t.u.  per  brake  horse  power  per  minute.  This  quantity 
will  increase  or  decrease  with  change  in  operating  conditions; 
its  limits  are  apparently  between  30  and  60  B.t.u.  per  brake 
horse  power  per  minute. 

The  cooling  surface  of  the  modem  air-cooled  cylinder  almost 
invariably  takes  the  form  of  a  series  of  fins.  Tests  on  the  rate 
of  heat  dissipation  from  such  surfaces,  made  by  the  British 
Advisory  Committee  for  Aeronautics,1  show  that,  for  wind 
speeds  between  20  and  60  miles  per  hour,  the  heat  loss  for  a 
given  material  is  independent  of  the  roughness  of  the 
surface.  Steel  shows  5  to  10  per  cent  greater  heat  dissipation 
than  aluminum  or  copper.  Aluminum  is  improved  about  10 
per  cent  by  a  coating  of  stove  enamel. 

Throughout  the  usual  range  of  cylinder  diameters  the  heat 
dissipation  for  copper  fins  in  a  parallel  air  blast  at  the  ground  is 
given  by 

H  =  [0.0247  -  0.0054(i°--.V^-4^0'78 

where  H  is  the  heat  dissipated  in  B.t.u.  per  square  foot  of  fin 
surface  per  minute  per  degree  Fahrenheit  difference  between  the 
mean  fin  temperature  and  the  incoming  air  temperature;  I  is  the 
length  of  the  fins  in  inches;  p  is  the  pitch  in  inches  measured  from 
surface  to  surface  of  adjacent  fins;  and  V  is  the  wind  speed  in 
miles  per  hour.  With  tapering  fins  p  should  be  taken  at  the 
mean  height  of  the  fin.  The  heat  dissipation,  depending  on  the 
weight  of  air  brought  into  contact  with  the  cylinder,  is 
proportional  to  (dF)0*73,  where  d  is  the  air  density. 

Shape  and  Size  of  Fins. — The  fin  which  gives  the  maximum 
heat-loss  per  unit  of  weight  is  one  having  slightly  concave 
surfaces  and  a  sharp  tip;  a  plain  triangular  fin  is  only  very  slightly 
less  efficient.  The  best  proportions  for  such  a  fin  depend  on  the 
conductivity  of  the  material  and  on  the  wind  speed.  For  a 
speed  of  40  miles  per  hour,  the  following  table  shows  the  best 

1  A.  H.  GIBSON,  Institution  of  Automobile  Engineers  of  Great  Britain,  1920. 


346 


THE  AIRPLANE  ENGINE 


proportions  for  fins  of  aluminum  alloy  (conductivity  =  0.38 
C.G.S.  units)  and  steel  (conductivity  =  0.12  C.G.S.  units)  and 
also  for  rectangular  copper  fins  (conductivity  =  0.90  C.G.S. 
units). 


Bottom  breadth,  centimeters 

0.025 

0.05 

0.1 

0.2 

0.3 

0.4 

0.5 

Length,  centimeters: 
Aluminum                    .  .  . 

2.0 

2.9 

3.5 

4.1 

4.5 

Steel     

1.1 

1.5 

1.8 

2.1 

?,  3 

Copper                             .  .  . 

1.6 

2.3 

3.3 

4.8 

If  such  a  fin  be  truncated  until  the  tip  breadth  is  one-fifth  of 
the  bottom  breadth,  the  lengths  become  80  per  cent  of  those 
given  above.  The  heat  dissipation  is  about  0.88  time  and  the 
weight  0.96  time  as  great  as 'for  the  complete  triangular  fin. 

Since  the  heat  dissipated  from  a  fin  of  given  shape  varies 
directly  as  the  length  of  the  fin,  while  the  weight  varies  as  the 
square  of  the  length  of  the  fin  (other  things  being  equal)  cooling 
fins  should  be  as  short  as  possible,  a  large  number  of  short  thin 
fins  being  used  in  preference  to  a  smaller  number  of  longer  and 
thicker  fins.  While  in  practice  this  is  to  be  borne  in  mind,  many 
other  factors  besides  that  of  weight  have  an  important  bearing 
on  the  best  size  of  fin  to  be  adopted.  Thus,  in  a  thin  steel 
cylinder,  or  in  a  cylinder  of  cast  iron  or  cast-aluminum  alloy,  the 
circumferential  ribs  add  greatly  to  the  strength  and  resistance  to 
distortion.  Comparatively  deep  and  heavy  fins  have  a  greater 
effect  in  this  direction  than  a  larger  number  of  similar  but  smaller 
fins  giving  the  same  cooling.  Again,  as  the  number  of  fins  in 
increased,  the  pitch  is  correspondingly  diminished.  This 
diminution  in  pitch  reduces  the  air  flow  between  the  fins  to  an 
extent  which  may,  with  very  small  pitches,  render  the  fins  prac- 
tically useless  for  cooling  purposes. 

In  a  cylinder  of  cast  iron  or  of  aluminum  alloy,  foundry 
difficulties  put  a  definite  limit  to  the  minimum  pitch  of  the  fins. 
On  the  barrel  itself  a  somewhat  smaller  pitch  may  be  adopted 
than  on  the  head,'  or  the  barrel  fins  may  be  turned  out  of  the 
solid  if  desired.  On  account  of  the  complicated  form  of  the  cylin- 
der head  and  ports,  however,  it  is  difficult  to  machine  their 
cooling  fins,  and  the  length  of  many  of  the  cores  necessitates  the 


THE  COOLING  SYSTEM 


347 


pitch  being  made  fairly  large.  The  minimum  practical  pitch 
of  fin  for  such  cylinders,  having  a  diameter  of  from  4  in.  to  6  in., 
is  about  8  to  9  mm.  or  about  <^LG  m-  Foundry  difficulties  also 
prevent  the  casting  of  a  fin  having  a  tip  less  than  about  0.5  mm. 
in  thickness,  or  a  root  thickness  less  than  about  Z/10,  so  that  an 
aluminum  fin  1  in.  long  would  not  have  a  root  thickness  less  than 
0.1  in. 

For  steel  cylinders  with  fins  turned  out  of  the  solid,  the  pitch 
may  with  advantage  be  cut  down  to  about  }£  in.  on  a  cylinder  of 
3  in.  or  so  in  diameter,  but  there  appears  to  be  little  to  be  gained 
by  reducing  the  pitch  beyond  this  point. 

The  mean  fin  temperature  depends  on  the  total  amount  of  heat 
which  has  to  be  dissipated  and  its  value  depends  on  many  factors. 
Determinations  of  the  actual  temperatures  of  the  cylinder 
walls,  pistons,  and  exhaust  valves  show  no  definite  differences 
between  well  designed  air-cooled  and  water-cooled  engines;  the 
air-cooled  cylinder  may  be,  and  often  is,  the  cooler  of  the  two. 

The  influence  of  engine  speed  on  the  wall  temperature  is 
shown  in  the  following  table  giving  the  temperature  at  a  point  on 
the  side  of  the  combustion  space  of  an  aluminum  air-cooled 
cylinder  operating  at  maximum  load. 


R.D.m 

800 

1,000 

a  ,  200 

1,400 

1,600 

1,800 

B.h.p 

10  2 

12  8 

15  4 

18 

19  7 

20  6 

Temperature,  degrees  Cen- 
tigrade 

100 

103 

124 

123 

136 

138 

The  compression  ratio  is  more  important  than  engine  speed 
in  determining  the  wall  temperature.  There  is  usually  a  definite 
compression  ratio  giving  minimum  wall  temperature;  variations 
on  either  side  increase  that  temperature.  The  following  table 
gives  tests  of  a  100  by  140  mm.  aluminum  air-cooled  cylinder 
with  varying  compression  ratio.  The  brake  mean  effective 
pressure  and  fuel  consumption  of  this  engine  are  noteworthy. 


Compression  ratio             

4  6 

5.0 

5.4 

5.8 

6.2 

6  4 

Brake  mean  effective  pressure,  Ib.  per 
sq.  in  

116.1 

119.3 

122.0 

125.0 

129.0 

123.0 

Fuel,  pounds  per  brake  horse  power  per 

0  530 

0  507 

0  490 

0  475 

0.480 

0  520 

Mean    barrel  temperature,  f  Top 
degrees  Centigrade  \  Bottom 

180 
105 

170 
95 

157 
89 

154 

85 

183 
110 

212 
135 

348 


THE  AIRPLANE  ENGINE 


The  increased  temperatures  in  the  last  two  columns  result  from 
preignitions  which  were  occasional  with  compression  of  6.2  and 
frequent  with  6.4. 

Increase  in  cylinder  diameter  diminishes  slightly  the  heat  loss 
to  the  walls  per  brake  horse  power;  the  experimental  evidence 
suggests  a  decrease  of  about  3.5  per  cent  for  10  per  cent  increase 
in  cylinder  diameter.  As  the  ratio  of  cooling  surface  to  b.h.p. 
varies  inversely  as  the  diameter  in  similar  cylinders,  the  ratio  of 
cooling  area  to  heat  given  to  the  walls  decreases  as  the  diameter 
increases.  The  temperature  difference  between  an  air-cooled 
cylinder  and  the  cooling  air  in  a  given  wind  may  be  taken  as 
inversely  proportional  to  D°-6,  where  D  is  the  cylinder  diameter. 

The  air-fuel  ratio  has  considerable  influence  on  the  heat 
transmitted  to  the  cylinder  walls.  The  cylinder  is  hottest 
with  the  weakest  mixture  capable  of  sustaining  maximum  load; 
or,  approximately,  with  an  air-fuel  ratio  of  13.5.  Further 
weakening  of  the  mixture  makes  a  cooler  cylinder  on  account  of 
the  reduction  in  brake  horse  power  and  in  the  heat  loss  per 
b.h.p.  At  the  same  time  it  gives  a  hotter  exhaust  valve.  The 
last  point  is  brought  out  in  the  following  table  giving  test 
results  for  a  100  by  140  mm.  air-cooled  aluminum  cylinder. 


Air-fuel  ratio  

11  1 

11  9 

13.8 

15.2 

15.7 

Brake  m.e.p.,  Ib.  per  sq.  in.  ... 
Fuel,   pounds  per  brake  horse 
power  per  hour 

122 
0  622 

122 
0  589 

119 
0  515 

116 
0  480 

114 
0.470 

Exhaust  valve  temperature,  de- 
grees Centigrade  

706 

717 

747 

752 

747 

An  increase  in  mixture  strength  beyond  that  necessary  for 
maximum  power  reduces  the  temperature  of  the  cylinder  sur- 
faces, as  shown  below  for  an  engine  operating  at  full  throttle  and 
constant  speed. 


Air-fuel  ratio  . 

10  5 

11  4 

12  9 

13.5 

15.4 

Cylinder  head   temperature,   degrees 
Centigrade  .  . 

200 

215 

237 

229 

215 

Tests  show  that  the  maximum  temperature  of  the  head  of  an 
air-cooled  cylinder  must  not  exceed  270°C.  for  satisfactory  work- 
ing. If  the  temperature  exceeds  280°  there  is  usually  trouble 


THE  COOLING  SYSTEM  349 

from  preignition.  Higher  working  temperatures  are  permissible 
with  larger  cylinders.  If  the  temperature  is  kept  at  200°  to 
220°C.,  the  economy  and  capacity  obtained  are  quite  as  good  as 
for  water-cooled  cylinders  of  similar  design  and  size. 

The  temperature  of  the  exhaust  valve  at  its  hottest  point 
should  not  exceed  720°C.;  with  valves  not  exceeding  1.5  in.  in 
diameter  it  is  possible  to  reduce  this  temperature  to  650°C. 

In  a  well  designed  aluminum  cylinder  of  the  overhead-valve  type, 
operating  in  a  60-mile-per-hour  wind,  a  provision  of  0.28  to  0.35 
sq.  ft.  of  cooling  surface  per  brake  horse  power  is  sufficient  to  give 
satisfactory  operation,  the  larger  area  applying  to  cylinders  of 
about  4-in.  bore,  and  the  smaller  to  cylinders  of  about  6-in.  bore. 
For  steel  or  cast-iron  cylinders  with  overhead  valves  this  area 
must  be  increased  about  50  per  cent  and  for  L-head  cast-iron 
cylinders  by  100  per  cent. 

At  reduced  wind  speeds  the  cylinder  temperature  increases; 
the  mean  temperature  difference  between  the  fins  and  the  air 
varies  inversely  as  T70"*.  Thus  in  a  given  series  of  tests  a  reduc- 
tion of  wind  speed  from  80  to  40  miles  per  hour  increased  the 
cylinder  temperature  from  229°  to  296°C.  There  are  practical 
difficulties  in  the  way  of  providing  sufficient  cooling  surface  for 
operation  at  full  throttle  below  certain  limiting  wind  speeds. 
The  wind  speeds  can  be  less  with  smaller  cylinders.  The  mini- 
mum air  velocity  for  good  performance  of  air-cooled  cylinders 
of  good  design  and  material  under  full  throttle  is  given  below. 
At  lower  air  speeds  partial  throttle  only  should  be  used. 

Diameter,  inches 23468 

Minimum  air  velocity,  miles  per  hour ...     30        40        50        70        90 

Cylinder  Materials. — With  cylinders  of  normal  design  the 
middle  portion  of  the  head  is  the  hottest  point.  Free  air  flow 
to  this  point  is  impeded  by  the  valve  ports  and  gears  so  that  it  is 
almost  impossible  to  provide  adequate  cooling  surface  there. 
The  heat  has  to  travel  outward  and  is  dissipated  mainly  from 
the  cooling  surface  surrounding  the  combustion  head.  It 
is  therefore  important  to  use  a  material  of  maximum  thermal 
conductivity.  The  three  practical  materials  for  cylinder  con- 
struction, steel,  cast-iron  and  aluminum  alloy,  have  conductivities 
(in  C.G.S.  units)  of  0.12,  0.10  and  0.38  respectively.  Aluminum 
is  consequently  the  most  desirable  material.  The  alloys  most 
suitable  for  cylinders  are  copper-aluminum  alloys  with  about 


350 


THE  AIRPLANE  ENGINE 


90  per  cent  of  aluminum.  The  high-zinc  alloys  are  unsuitable 
because  their  tensile  strength  is  low  at  200°C.  All  the  alloys 
show  rapid  decrease  of  strength  as  the  temperature  increases 
beyond  250°C.  The  following  table  gives  data  on  this  point. 


Composition  per  cent 

Tensile  strength, 
Ib.  per  sq.  in. 

Cu 

Sn 

Mg 

Al 

Ni 

Mn 

At  250°C. 

At  350°C. 

7.0 

1.0 

92.0 

12,300 

6,700 

12.0 

88.0 

, 

.  . 

23,500 

13,400 

14.0 

85.0 

.  . 

1 

21,300 

14,500 

4.0 

1.5 

92.5 

2 

.  . 

24,600 

11;200 

9.0 

89.0 

2 

19,000 

10,100 

Water  Cooling. — By  water-cooling  the  cylinder  and  the 
exhaust  ports  it  is  possible  to  run  with  higher  speeds  and  com- 
pression ratios  than  are  practicable  with  air-cooled  cylinders. 
The  possible  increase  in  speed  and  ratio  of  compression  are  rel- 
atively unimportant  when  compared  with  the  performance  of 
the  best  recent  constructions  in  air-cooled  cylinders;  they  are 
considerable  as  compared  with  the  average  air-cooled  cylinder. 

With  water  cooling  it  is  possible  to  maintain  almost  any 
desired  cylinder  temperature.  If  the  temperature  is  low  the 
volumetric  efficiency  and  the  capacity  of  the  engine  will  be 
improved  (see  p.  37)  but  the  engine  friction  increases  and  its 
efficiency  falls  off.  The  temperature  of  the  jacket  water  after 
leaving  the  radiator  must  be  below  the  boiling  point  of  water  at 
the  pressure  existing  on  the  suction  side  of  the  pump,  otherwise 
the  pump  will  not  function  well  but  will  suck  in  water  vapor. 
As  fuel  economy  is  ordinarily  more  important  than  capacity, 
the  jacket  water  is  usually  kept  at  as  high  a  temperature  as 
the  boiling  point  will  permit.  The  mean  jacket  temperature 
at  the  ground  is  usually  160  to  180°F. 

Water  is  not  the  ideal  cooling  agent.  A  less  volatile  fluid  would 
permit  higher  cylinder  temperature;  higher  efficiencies  might  be 
obtained  without  running  into  such  temperatures  as  would  cause 
preignition.  The  same  result  might  be  obtained  by  operating  a 
closed  water-cooling  system  under  pressures  greater  than 
atmospheric,  but  this  would  necessitate  heavier  material  for  the 


THE  COOLING  SYSTEM 


351 


radiator  core  and  consequent  increase  in  weight  and  decrease  in 
airplane  efficiency.  The  heat  which  is  removed  by  the  jacket 
water  is  practically  equal  to  the  b.h.p.  or  is  42.4  B.t.u.  per  brake 
horse  power  per  minute.  In  a  closely  cowled  engine  this  same 
amount  of  heat  would  have  to  be  removed  from  the  radiator. 
With  the  usual  cowling  there  is  considerable  removal  of  heat 
by  the  air  stream  from  the  engine  and  water-j  acket  surfaces,  so 
that  only  about  31  B.t.u.  per  brake  horse  power  per  minute  has 
to  be  removed  from  the  radiator:  with  an  uncowled  engine  this 
quality  falls  to  23  or  25  B.t.u. 

The  principal  parts  of  a  water-cooling  system  are  the  jackets, 
the  pump  and  the  radiator.  The  last  of  these  will  be  considered 
first. 

Radiators. — Airplane  radiators  have  developed  from  auto- 
mobile practice  but  certain  types  of  automobile  radiators  are 
entirely  unsuited  to  air- 
plane practice.  The  suc- 
cessful commercial  types 
have  cores  made  of  thin 
brass,  or  copper  ribbons 
or  tubes  from  0.004  to 
0.006  in.  thick.  Common 
types  are  shown  in  Fig. 
269,  which  illustrates:  a 
and  6,  rectangular  air  pas- 
sages; Cj  rhombic  passages; 
and  d,  circular  passages  with 
hexagonal  ends.  Other  com- 
mon types  have  hexagonal 
or  elliptical  air  passages 
The  water  passages  are  nar- 
row, varying  from  0.03  to 
0.08  in.  The  air  tubes  are  commonly  not  more  than  Y±  in. 
in  maximum  cross-section  dimension  and  are  from  3  to  5  in. 
long  (depth  of  core).  The  metal  sheets  are  stamped  or  rolled 
to  the  desired  form  with  the  front  and  rear  ends  of  the  pair  of 
sheets  forming  each  water  passage  in  contact  with  one  another. 
These  ends  are  soldered  by  dipping  them  into  a  shallow  pool  of 
molten  solder.  Great  care  must  be  exercised  to  keep  down  the 
weight  of  solder  as  much  as  possible;  it  often  amounts  to  25  per 
cent  of  the  total  weight  of  the  radiator  core.  The  top  and 


FIG.  269. — Types  of  radiator  core. 


352  THE  AIRPLANE  ENGINE 

bottom  ends  of  the  water  passages  are  inserted  through  slots  in 
the  top  and  bottom  headers  respectively;  the  two  sheets  of  each 
water  passage  are  spread  apart  and  soldered  to  the  header. 
In  the  case  of  type  d,  Fig.  269,  the  ends  of  the  circular  tubes  are 
expanded  into  hexagonal  forms  which  are  soldered  together;  the 
expansion  is  made  enough  to  give  the  desired  width  of  water 
passage  between  the  tubes.  Type  b  differs  from  type  a  not 
only  in  the  method  of  assembly  but  may  also  be  made  of  a  cor- 
rugated surface  which  is  intended  to  give  greater  strength  and 
larger  radiating  surface.  The  types  a,  6,  c  and  d  in  Fig.  269  have 
water  in  contact  with  all  the  radiating  surface  and  are  said  to 
have  only  "direct"  radiating  surface.  Many  automobile 
radiators  have  extensions  of  this  direct  radiating  surface  in  the 
form  of  fins  on  flat  or  circular  tubes  (e,  Fig.  269),  metal  spirals, 
and  so  forth.  Such  " indirect"  radiating  surface  is  found  to 
have  too  high  a  ratio  of  head  resistance  to  heat-removing  capacity 
tO;be  satisfactory  for  airplane  use. 

The  dimensions  or  external  shape  of  a  radiator  can  be  adapted 
to  suit  its  location  and  desired  performance.  The  location 
may  be  such  that  air  may  pass  through  or  around  it  without 
obstruction,  in  which  case  it  is  said  to  be  in  an  " unobstructed " 
position.  On  the  other  hand,  the  radiator  may  be  located  in  the 
nose  of  the  fuselage,  or  in  the  plane  of  the  wing,  in  which  case 
the  air  flow  is  materially  affected  by  other  parts  of  the  plane 
and  the  radiator  is  said  to  be  "  obstructed."  The  performance  of 
such  a  radiator  will  depend  not  only  on  the  size  and  type  of  the 
core  but  on  its  position  or  surroundings.  Examples  of  typical 
unobstructed  locations  are  shown  in  Fig.  270  (at  the  sides  of 
fuselage)  and  in  Fig.  271  (over  the  engine);  common  obstructed 
positions  are  in  the  nose  of  the  fuselage,  and  in  the  wing. 

A  comprehensive  study  of  the  properties  of  various  types  and 
dimensions  of  radiator  cores  has  been  made  at  the  Bureau  of 
Standards  and  published  in  the  Fifth  Annual  Report  of  the 
National  Advisory  Committee  on  Aeronautics.  The  following 
discussion  is  mainly  from  that  source. 

Two  quantities  are  of  importance  in  determining  the  heat 
transfer  of  a  core.  They  are  the  temperature  difference  between 
the  entering  air  and  the  mean  water  temperature;  and  the 
mass  flow  of  air.  The  temperature  difference  should  ordinarily 
be  taken  as  the  difference  between  the  mean  summer  air  tempera- 
ture and  the  mean  water  temperature.  The  mass  flow  of  air, 


THE  COOLING  SYSTEM 


353 


Outlet  Pipe. 


FIG.  270. — Side  radiators. 


FIG.  271. — Overhead  radiators. 


23 


354  THE  AIRPLANE  ENGINE 

M,  is  the  weight  of  air  flowing  per  second  per  square  foot  of 
frontal  area  of  the  core.  Its  amount  (at  constant  air  density) 
is  found  to  be  proportional  to  the  free  air  speed  or  the  velocity 
with  which  the  core  moves  through  the  air  when  the  core  is 
unobstructed;  the  mass  flow  is  always  less  for  obstructed  positions 
than  for  unobstructed. 

The  energy  dissipated  or  heat  transfer  is  expressed  in  horse 
power  per  square  foot  of  frontal  area,  and,  for  purposes  of  com- 
parison of  the  properties  of  various  cores,  a  temperature  difference 
of  100°F.  is  assumed  between  the  air  entering  the  radiator  and  the 
mean  water  temperature;  the  heat  transfer  is  proportional  to  this 
temperature  difference.  One  horse  power  is  equivalent  to  42.54 
B.t.u.  per  minute. 

The  head  resistance  of  the  core  is  the  force  required  to  push  it 
through  the  air  and  is  expressed  in  pounds  per  square  foot  of 
frontal  area.  This  head  resistance  is  found  to  vary  approxi- 
mately as  the  square  of  the  free  air  speed;  in  most  cases  the 
exponent  is  slightly  less  than  2.  If  R  is  the  head  resistance, 
and  V  the  free  air  speed  in  miles  per  hour,  then 

R  =  cV2 

and  c  is  called  the  head  resistance  constant. 

The  horse  power  absorbed  by  a  radiator  is  the  engine  power 
required  to  overcome  the  head  resistance  and  support  the  weight 
of  the  radiator.  The  work  done  in  supporting  the  weight  can  be 
calculated  if  the  lift-drag  ratio  of  the  plane  as  a  whole  is  known. 
An  average  value  of  5.4  may  be  assumed  for  this  ratio.  If  W 
is  the  weight  of  the  core  and  contairied  water  in  pounds  per 
square  foot  of  frontal  area,  the  propeller  thrust  required  to  sup- 
port the  weight  is  TF/5.4.  The  horse  power  absorbed  is 

V  X  5'280 
'  60X~3pOO 


It  should  be  noted  that  this  method  of  calculation  neglects  the 
effect  on  the  lift-drag  ratio  of  the  addition  of  the  radiator. 
The  lift-drag  ratio  varies  between  different  planes  and  varies 
even  more  widely  between  climbing  and  level  flight.  The  selec- 
tion of  a  core  for  a  given  plane  cannot  be  made  satisfactorily 
without  a  knowledge  of  the  relative  importance  of  climbing 
speed  and  top  speed.  A  lift-drag  value  of  5.4  is  a  good  average 


THE  COOLING  SYSTEM  355 

and  gives  about  equal  value  to  climbing  and  level  speed.  If  the 
rate  of  climb  is  of  prime  importance  the  value  may  be  as  low  as 
3,  while  if  speed  on  the  level  is  the  most  important,  a  value  as 
high  as  10  may  be  used. 

A  small  additional  power  charge  against  the  radiator  is  that 
required  to  overcome  the  resistance  to  water  circulation  in  the 
radiator.  It  is  usually  so  small  as  to  be  negligible. 

The  definition  just  given  of  horse  power  absorbed  applies  only 
to  the  case  of  an  unobstructed  radiator.  If  the  addition  of  a 
radiator  necessitates  alterations  in  structure  (such  as  the  sub- 
stitution of  a  flat  nose  for  a  stream-line  fuselage  or  the  enlarge- 
ment of  the  fuselage  to  accommodate  the  radiator  required) 
the  consequent  increase  in  resistance  of  the  structure  should  be 
charged  to  the  radiator. 

A  comparison  of  the  performance  of  various  cores  can  be 
obtained  when  the  heat  transfer  per  unit  of  power  absorbed 
is  known.  This  quantity  is  called  the  Figure  of  Merit  and  is 
a  pure  number.  The  comparison  must  be  for  the  same  tempera- 
ture difference  and  free  air  speed.  It  applies  only  to  unob- 
structed radiators. 

The  general  conclusions  derived  from  the  tests  at  the  Bureau 
of  Standards  are  as  follows: 

Heat  transfer  is  a  function  of  mass  flow  of  air  and  is  inde- 
pendent of  the  air  density. 

Heat  transfer  is  roughly  proportional  to  mass  flow  for  a  core 
having  only  direct  cooling  surface.  When  there  is  a  considerable 
amount  of  indirect  cooling  surface  the  heat  transfer  increases  less 
rapidly  than  mass  flow  at  high  air  speeds. 

Heat  transfer  is  proportional  to  the  temperature  difference. 
'[Heat  transfer  is  not  greatly  affected  by  the  rate  of  water  flow  pro- 
vided the  rate  is  above  2  gal.  per  minute  per  inch  of  core  depth  per 
foot  width  of  core.  It  should  be  noted,  however,  that  this  is  true 
only  when  the  mean  water  temperature  is  regarded  as  constant. 

Heat  transfer  from  direct  cooling  surface  is  not  appreciably 
affected  by  the  composition  of  the  metal.  When  fins  and  other 
indirect  cooling  surface  are  used  the  thermal  conductivity  of  the 
metal  is  important. 

Heat  transfer  is  somewhat  increased,  but  at  the  expense  of  a 
large  increase  in  head  resistance,  by  spirals  or  other  forms  of 
passages  which  increase  the  turbulence  of  the  air  stream.  Heat 
transfer  is  greater  for  smooth  than  for  rough  tube  walls,  for,  if 


356  THE  AIRPLANE  ENGINE 

the  surface  is  rough,  it  will  be  covered  with  a  layer  of  more  or  less 
stagnant  fluid. 

Head  resistance  for  any  particular  core  varies  approximately 
as  the  square  of  the  free  air  speed. 

The  head  resistance  of  a  core  appears  to  be  closely  related  to 
its  mass  flow  so  that,  in  general,  anything  which  tends  to  cut 
down  the  flow  of  air  through  the  core  will  cause  a  considerable 
increase  in  head  resistance. 

Head  resistance  varies  directly  as  the  air  density  for  a  given 
free  air  speed,  and  inversely  as  the  density  for  a  given  mass  flow. 

Head  resistance  is  considerably  increased  by  projections, 
indentations,  or  holes  in  the  air  tube  walls. 

Head  resistance  per  square  foot  is  not  appreciably  affected  by 
the  size  of  the  core  within  the  limits  used,  viz.,  8  by  8  in.  to  16  by 
16  in.  and  12  by  24  in. 

Special  conclusions  with  reference  to  types  of  cores  are  as 
follows : 

For  a  high  figure  of  merit  the  core  should  have  smooth,  straight 
air  passages,  easy  entrances  and  exits  for  the  air  and  a  large  per- 
centage of  free  area.  Under  these  conditions  the  figure  of  merit 
increases  as  the  depth  increases  up  to  at  least  20  times  the  di- 
ameter of  the  air  tubes,  which  is  as  far  as  experiment  has  gone. 
Even  greater  depths  may  be  of  advantage. 

By  far  the  most  satisfactory  radiator  for  use  in  unobstructed 
positions  seems  to  be  one  of  thin  flat  plates  with  water  space  not 
over  KG  in.  wide  and  spaced  J^  in.  on  centers.  The  plate  should 
be  at  least  12  in.  deep.  As  the  figure  of  merit  changes  but 
slightly  with  increase  of  depth  beyond  12  in.  the  depth  may  be 
made  20  in.  or  more  if  it  is  desirable  to  reduce  frontal  area.  The 
chief  defect  of  the  type  is  mechanical  weakness.  Of  the  commer- 
cial radiators  tested,  those  have  given  highest  figure  of  merit  at 
high  air  speeds  which  have  only  direct  cooling  surface  in  the 
form  of  tubes  about  J^  m-  square  and  about  5  in.  deep.  The 
figure  of  merit  of  this  type  at  120  miles  per  hour  free  air  speed 
varies  from  about  8  to  8.4,  whereas  flat  plates  9%  in.  deep  and 
J^  in.  on  centers  have  a  value  of  10.7.  The  energy  dissipated  per 
square  foot  of  frontal  area  is  less  in  the  above  flat  plate  radiator 
than  in  the  best  square  tube  radiators  so  that  a  larger  frontal 
area  will  be  required  with  flat  plates  but  the  power  absorbed 
will  be  less. 

The  British  Air  Ministry  has  adopted  as  standard  a  circular 


THE  COOLING  SYSTEM  357 

tube  10  mm.  in  diameter  expanded  at  the  ends  to  a  hexagonal 
section  11  mm.  across  the  flats  (Fig.  269  d).  The  standard  length 
of  the  tubes  is  120  mm.  The  material  is  70  —  30  brass  with  wall 
thickness  of  0.005  in. 

The  actual  power  absorbed  by  the  radiator  in  being  lifted  and 
pushed  through  the  air  (see  Table  18)  varies  from  about  3  h.p. 
to  6  h.p.  per  square  foot  of  frontal  area  at  100  miles  per  hour. 
This  amounts  to  from  5  to  20  per  cent  of  the  total  engine  power. 
A  small  gain  in  radiator  performance  may  have  an  appreciable 
effect  at  high  speed. 

The  selection  of  a  radiator  core  for  an  obstructed  position  is 
more  difficult.  An  obstructed  position  involves  a  large  absorp- 
tion of  power.  The  resistance  of  a  fuselage  fitted  with  a  nose 
radiator  is  two  or  three  times  the  resistance  of  the  same  fuselage 
with  a  stream-line  nose.  The  increase  in  resistance  due  to  the 
substitution  of  a  radiator  for  a  stream-line  nose  is  greater  than 
the  increase  that  would  be  caused  by  using  a  radiator  of  the 
same  core  construction  and  the  same  cooling  capacity  in  an 
unobstructed  position. 

At  any  given  free-air  speed  the  total  resistance  of  a  fuselage 
with  a  flat  nose  radiator  is  increased  by  increasing  the  air  flow 
through  the  radiator,  either  by  opening  exit  vents  for  the  air  or 
by  decreasing  the  resistance  of  the  radiator  to  the  passage  of  air. 
This  indicates  that  a  nose  radiator  should  be  of  compact  con- 
struction with  high  heat  transfer,  for  low  air  flows  through  the 
core,  requiring  a  core  of  high  resistance.  This  fact  is  of  special 
importance  since  the  space  available  for  a  nose  radiator  is  so 
limited  that  the  highest  possible  mass  flows  are  used  in  practice. 
A  nose  radiator  with  air  exit  vents  equal  in  area  to  the  free  air 
passage  through  the  radiator  is  found  to  cut  down  the  heat 
transfer  about  35  per  cent  as  compared  with  the  same  radiator 
in  an  unobstructed  position.  Indirect  cooling  surface  may  be  of 
advantage  if  it  is  made  of  copper,  crimped  from  the  water  tube 
walls  and  well  soldered  to  them  at  all  possible  places.  Several 
types  of  core  show  good  heat  transfer  at  low  speeds,  but  here 
again  the  square  tube,  with  direct  radiating  surface  only,  gives 
best  result  of  all  commercial  types,  and  flat  plates  spaced  J4  m- 
on  centers  show  excellent  performance.  The  fin  and  tube  type 
with  its  small  amount  of  direct  surface  has  no  use  in  airplanes 
except  possibly  in  a  wing  position  where  a  high  head  resistance 
is  no  disadvantage. 


358 


THE  AIRPLANE  ENGINE 


The  properties  of  cores  selected  as  typical  of  common  con- 
struction are  given  in  Tables  17  and  18.  The  constructions  vary 
from' the  flat  tubes  (E-Q  and  E-8)  with  100  per  cent  direct  cooling 
surface  and  a  minimum  of  head  resistance  to  finned  circular 
tubes  (F-5)  with  only  12.3  per  cent  of  direct  surface  and  three 
times  as  much  head  resistance  per  square  foot  of  frontal  area  as 
the  best  flat  tubes.  The  dimensions  of  the  core  are  given  in  Table 
17;  the  performance  at  various  wind  speeds  in  unobstructed  posi- 
tions in  Table  18.  The  "  figure  of  merit "  necessarily  diminishes 


\ 


0          0.020      0.040        0.060       0.050       0.100 
Mass  Flow  Factor  (k) 

FIG.  272. — Head  resistance  constant  and  mass  flow  factor. 


with  increase  of  wind  speed;  the  order  of  merit  of  the  different 
cores  is  different  at  different  speeds.  At  120  miles  per  hour  the 
flat  tubes  E-8  and  E-Q  are  seen  to  be  best  and  the  finned  circular 
tubes  F-5  the  poorest. 

Table  19  gives  the  constants  in  the  empirical  equations  for 
"Head  Resistance,"  "Mass  Flow"  and  "Energy  Dissipated"  for 
the  cores  listed  in  Table  17;  these  quantities  are  obtained  from 
the  data  of  Table  18.  The  Head  Resistance  Constant  and  the 
Mass  Flow  Constant  appear  to  be  connected  by  a  simple  relation; 
plotting  these  quantities  for  all  the  cores  tested  gives  the  curve  of 
Fig.  272. 


THE  COOLING  SYSTEM 


359 


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360 


THE  AIRPLANE  ENGINE 


TABLE  18. — RADIATOR  PERFORMANCE  IN  TERMS  OF  FREE  AIR  SPEED  (FOR 

UNOBSTRUCTED  POSITIONS  ONLY) 
Grade  A  represents  very  good  performance;  grade  E,  very  poor 


Radiator 

Speed, 
miles 
per 
hour 

Air  flow, 
Ib.  per 
sq.  ft. 
per  sec. 

Energy  in  h.p. 
dissipated  per 
square  foot  of 

Head 
resistance, 
Ib.  per 
sq.  ft. 
frontal 
area 

H.p. 
absorbed 
per  sq. 
ft.  of 
frontal 
area 

Figure 
of  meri 

Front 

Surface 

A-7....... 

J»" 
Grade  

30 
60 
90 
120 

2.20 
4.40 
6.60 
8.80 

27.2 
45.9 
61.7 
76.5 

A 

20.3 
37.2 
52.9 
68.0 

B 
20.2 
31.0 

39.8 
47.9 

D 

14.8 
26.4 
37.8 
48.5 

D 
13.8 
24.7 
32.3 
38.7 

D 

29.3 
51.1 
71.3 
90.5 

A 
17.2 
29.8 
41.0 
51.7 
C 
13.8 
21.7 
27.8 
33.0 

E 
23.1 
39.5 
53.6 
67.0 

B 

0.43 
0.73 
0.99 
1.22 

0.37 
0.68 
0.97 
1.25 

0.36 
0.55 
0.71 
0.85 

0.46 
0.83 
1.18 
1.52 

0.58 
1.04 
1.32 
1.63 

0.37 
0.65 
0.91 
1.15 

0.44 
0.76 
1.05 
1.32 

0.40 
0.63 
0.81 
0.96 

0.52 

0.88 
1.20 
1.50 

1.72 
6.70 
14.90 
26.3 

D 
1.40 
5.61 
12.10 
21.3 

C 
1.72 
6.88 
15.48 
27.5 

D 
1.37 
5.47 
12.30 
21.9 

C 
1.25 
4.75 
10.48 
18.4 

B 

1.57 
6.27 
14.10 
25.1 

D 
0.78 
3.12 
7.03 
12.5 
A 
2.52 
9.65 
21.6 
38.3 

E 
2.40 
8.65 
19.1 
33.4 

E 

.50 
1.79 
4.66 
9.87 

D 
0.44 
1.54 

3.88 
8.11 

C 
0.41 
1.64 
4.52 
9.89 

D 
0.30 
1.26 
3.53 

7.79 

C 
0.26 
1.08 
2.99 
6.53 

B 
0.51 
1.78 
4.54 
9.59 

D 
0.27 
0.92 
2.31 
4.84 
A 
0.33 
1.81 
5.57 
12.8 

E 
0.40 
1.80 
5.21 
11.5 

E 

54.6 
25.7 
13.2 

7.8 

.     B 
46.6 
24.1 
13.6 

8.4 

B 
49.8 
18.9 
8.8 
4.8 

D 

48.7 
20.9 
10.7 
6.2 

C 
53.5 
23.0 
10.8 
5.9 

C 

57.2 
28.8 
15.7 
9.5 

B 
63.3 
32.4 
17.8 
10.7 
A 
41.3 
12.0 
5.0 
2.6 

E 
58.0 
22.0 
10.3 
5.8 

C 

A-23  

30 
60 
90 
120 

2.29 
4.58 
6.87 
9.16 

9$ 
Grade  

B-8  

30 
60 
90 
120 

2.12 
4.24 
6.36 
8.48 

Grade  

C-4     

30 
60 
90 
120 

2.40 

4.80 
7.20 
9.60 

Grade 

D-l  

30 
60 
90 
120 

30 
60 
90 
120 

2.41 
4.82 
7.22 
9.63 

Grade 

E-6   

2.12 
4.24 
6.36 
8.48 

-Hh&i~ 

Grade  

E-8 

30 
60 
90 
120 

2.74 
5.48 
8.23 
10.97 

1  " 
l^L 

Grade  

F-5 

30 
60 
90 
120 

1.82 
3.64 
5.46 
7.28 

-  -kl'd?     ©  © 

t|mX    HI" 

—  '  %'MJ:     \®  O 
i  i  i  1     IK**     .  H 

~~*^        100 

Grade  

G-3 

30 
60 
90 
120 

1.88 
3.75 
5.62 
7.50 

Grade  

THE  COOLING  SYSTEM 


361 


A  plotting  of  the  data  in  Table  18  for  core  E-S  is  given  in 
Fig.  273. 

Size  of  Radiator. — If  a  type  of  core  has  been  selected  and  its 
properties  are  known  and  plotted  as  in  Fig.  273,  the  necessary 
size  of  an  unobstructed  core  is  directly  obtainable.  The  dimen- 
sions must  be  calculated  for  the  most  unfavorable  condition, 
which,  ordinarily,  will  be  maximum  climbing  speed  near  the 
ground  with  summer  temperatures.  The  mean  water  temper- 
ature will  be  fixed  by  the  maximum  temperature  allowable  in 
the  jackets  and  by  the  quantity  of  water  circulated.  The 

Free  Air  Speed,  Mi.perHr. 

50  IQO 


39.2  Sq.H:  Surf  ace 
!00%  Direct  Surface 
975% free  Anta 
32  Ib.Fmpfy 


0246  8  10 

Mass,  Flow  of  Air,  Lb.  perSq.FtperSec. 
FIG.  273. — Properties  of  a  flat  tube  radiator  core. 

maximum  allowable  water  temperature  is  ordinarily  20  to  30°F. 
below  the  boiling  point  and  varies  with  the  altitude  of  the  plane; 
its  value  will  be  determined  by  (1)  its  influence  on  the  volumetric 
efficiency  of  the  engine  (see  p.  37)  and  (2)  the  water  resistance 
of  the  -radiator  (see  p.  364).  If  the  water  system  is  closed, 
with  no  vent  to  the  atmosphere,  the  last-mentioned  factor 
disappears.  The  heat  to  be  dissipated  should  be  determined  if 
possible  but  may  be  assumed  equal  to  the  brake  work  of  the 
engine  if  more  exact  knowledge  is  not  obtainable.  The  effect 
of  propeller  slip  should  be  estimated  and  allowed  for.  Allowance 
should  also  be  made  for  the  cooling  effect  of  the  radiator  headers 
and  of  the  exposure  of  the  engine  to  the  wind. 


362 


THE  AIRPLANE  ENGINE 


Occasionally,  instead  of  designing  for  maximum  climb  some 
other  condition  may  impose  maximum  service  on  the  radiator 
as,  for  example,  in  flying  boats  and  seaplanes  intended  for 
training,  where  much  taxi-ing  is  done  at  low  plane  speed  and 
maximum  engine  power.  If  the  radiator  is  in  the  nose  of  the 
fuselage  some  assumption  must  be  made  as  to  the  relation  of  mass 
flow  to  the  speed  of  the  plane.  The  mass  flow  will  usually  vary 
from  0.04  to  0.07  time  the  speed  of  the  plane  (in  miles  per  hour), 
depending  on  the  type  of  radiator,  the  amount  of  cooling  and  the 
masking  effect  of  the  propeller.  The  power  absorbed  is  seldom 
calculable  because  of  the  uncertain  effect  of  the  radiator  on  the 
resistance  of  the  fuselage. 

TABLE  19. — CONSTANTS  IN  THE  EQUATIONS  R  =  cV2;M  =/cF/ANDQ  =  Gmn 

R  =  Head  resistance  in  pounds  per  square  foot. 

V  =  Free-air  speed  in  miles  per  hour. 
M  =  Mass  flow  of  air  in  pounds  per  second  per  square  foot. 

Q  =  Energy  dissipated  in  horsepower  per  square  foot  per  100°F.  tem- 
perature difference. 

m  =  "mass  flow  constant,"  which  is  the  ratio  of  the  mass  of  air  passing 
through  1  square  foot  of  radiator  to  the  mass  of  air  passing  through 
1  square  foot  of  free  area  in  front  of  the  radiator. 


Radiator 

c  X  103 

k  X  102 

m 

G 

n 

A-7 

1  86 

7  34 

0  667 

15  1 

0  75 

A  -23  

1.56 

7  63 

0  694 

10  3 

0.85 

B-8 

1  91 

7  07 

0  643 

13  1 

0  60 

C-4  

1  52 

8.00 

0  727 

7  1 

0.85 

0-1, 

1  32 

8  03 

0  730 

8  1 

0  70 

E-Q  

1.74 

7.07 

0  643 

16.1 

0.80 

E-8. 

0  867 

9  13 

0  830 

7  6 

0  80 

F-5  

2.68 

6.07 

0.552 

10.0 

0.60 

G-3. 

2  40 

6  25 

0  568 

14  8 

0  75 

The  mass  flow  for  a  wing  radiator  depends  on  the  angle  of 
incidence  but  is  probably  not  over  0.01  time  the  plane  speed 
even  at  the  best  climbing  angle. 

The  relative  efficiencies  of  radiators  in  various  positions  are 
given  by  Liptrot1  as  follows : 

1  Aeronautics,  Apr.  29,  1920. 


THE  COOLING  SYSTEM  363 


Relative 
Position  of  radiator  ~,  . 

efficiency 


Unobstructed 

Underslung,  side  or  overhead,  but  close  to  fuselage 

Twin  nose  radiator 

Nose  radiator  with  core  entirely  above  or  below  propeller 


shaft . . 


Nose  radiator  with  propeller  shaft  in  center 

Behind  engine 


1.000 
0.973 
0.716 


0.656 
0.585 
0.423 


The  use  of  a  small  projecting  lip  or  stream  line  entrance  around 
the  core  may  reduce  the  necessary  core  size  slightly  but  at  the  cost 
of  a  considerable  increase  of  head  resistance. 

Rate  of  Water  Flow.  —  One  gallon  (231  cu.  in.)  of  water  at 


200°F.     weighs  -^-=  8   Ib.    approximately.     With    a 


temperature  difference  of   10°F.,    1   gal.   of  water  per  minute 

SO 
will  give  up  80  B.t.u.  or  ,0  A.  =  2  h.p.  approximately.     With  a 

4Z.4O 

temperature  difference  of  5°F.  the  flow  of  water  in  gallons  per 
minute  should  equal  the  engine  horse  power. 

The  entering  temperature  of  the  water  is  fixed  by  the  necessity 
of  keeping  at  a  certain  point  below  boiling.  With  fixed  entering 
temperature,  if  the  amount  of  water  circulated  is  increased  the 
mean  temperature  of  the  water  is  raised  and  consequently  the 
temperature  difference  between  air  and  water  is  increased.  The 
influence  of  water  velocity  on  the  heat  transfer  is  found  by 
experiment  to  be  very  small  so  long  as  the  velocity  is  above  2  gal. 
per  minute  per  foot  width  per  inch  depth  of  core,  which  is  much 
below  usual  rates.  With  a  circulation  of  J^  gal.  per  minute  per 
horse  power  the  temperature  fall  of  the  water  is  20°F.  ;  increasing 
this  to  %  gal.  reduces  the  temperature  fall  of  the  water  to  10°F. 
and  increases  the  temperature  difference  between  air  and  water 
by  5°F.  With  an  infinite  amount  of  water  circulated  this  tem- 
perature difference  could  be  increased  only  another  5°F.  The 
increase  in  pump  work  with  increased  water  flow  makes  it 
undesirable  to  circulate  more  than  about  %  gal.  per  minute  per 
horse  power,  and  with  radiators  that  are  relatively  long  and 
narrow  a  flow  of  ^  gal.  per  minute  per  horse  power  should  be 
used. 


364 


THE  AIRPLANE  ENGINE 


The  pressures  required  to  maintain  water  flow  through  the 
cores  of  radiators  vary  greatly  with  the  dimensions  and  type  of 
construction.  Those  types  having  the  widest  and  straightest 
water  spaces  offer  least  resistance  whereas  those  with  many  right 
angle  bends  will  offer  much  resistance.  The  range  for  12  com- 
mercial radiators  tested  at  the  Bureau  of  Standards,  all  of  them 
8  in.  square  in  frontal  section  and  of  depths  varying  from  2%  to 
4  in.  with  a  total  water  flow  of  20  gal.  per  minute,  was  from  0.27 
to  10.2  ft.  of  water  pressure  drop.  These  pressure  drops  may  be 
assumed  to  vary  directly  as  the  height  of  the  core,  but  the  rate  of 
change  with  change  of  water  velocity  follows  an  exponential  law 
in  ah1  cases,  though  with  a  widely  varying  exponent  in  the 
different  types.  The  resistance  seems  to  depend  largely  on  the 
care  used  in  manufacture  and  on  the  form  of  the  water  tube 
entrances  and  exits.  It  would  seem  well  to  include  a  test  for 
pressure  necessary  to  produce  water  flow  in  acceptance  specifi- 
cations for  complete  radiators. 

The  water  enters  the  top  header  of  the  radiator,  at  which  place 
atmospheric  pressure  is  usually  maintained  through  the  overflow 
pipe.  The  suction  pressure  at  the  pump  cannot  be  less  than  the 

vapor  pressure  of  the 


Tank  on  Top  Plane 


I 


'•  Carburetors  '  ,., ., 

Radiators  on  each  side 

of  Machine 
FIG.  274. — Cooling  system  of  Benz  engine. 


water  leaving  the  rad- 
iator if  the  pump  and 
radiator  are  at  the  same 
level.  If  the  water  leaves 
the  radiator  at  190°F. 
the  corresponding  vapor 
pressure  is  9.2  Ib.  or 
about  5  Ib.  below  at- 
mospheric pressure.  The 
maximum  pressure  avail- 
able for  overcoming  the 


resistance  of  the  radiator  in  this  case  will  be  5  Ib.  per  square  inch 
or  11.5  ft.  of  water.  With  a  reserve  tank  in  the  upper  plane,  as 
in  Fig.  274,  the  head  available  in  overcoming  radiator  friction 
is  increased  by  the  height  of  the  tank  above  the  suction.  If 
the  resistance  of  a  proposed  radiator  is  in  excess  of  the  available 
pressure,  its  height  must  be  decreased  and  its  width  correspond- 
ingly increased  in  order  to  give  the  necessary  radiating  surface. 

Occasionally  radiation  or  expansion  tanks  instead  of  being 
vented  to  the  atmosphere  are  provided  with  safety  valves  opening 


THE  COOLING  SYSTEM 


365 


at  2  or  3  Ib.  per  square  inch.     This  diminishes  the  loss  of  water 
from  evaporation  and  may  permit  a  higher  water  temperature. 

Effect  of  Altitude  on  Radiator  Performance. — The  investiga- 
tions at  the  Bureau  of  Standards  have  yielded  the  following 
general  conclusions: 

The  effect  of  the  lower  air  temperature  is  to  increase  the  heat 
transfer  in  proportion  to  the  increase  in  the  mean  temperature 
difference  between  the  entering  air  and  the  water.  The  decrease 
in  air  density  reduces  the  mass  flow  of  air  and  decreases  the  heat 
transfer  at  any  given  plane  speed  in  proportion  to  the  air  density. 

Head  resistance  is  proportional  to  air  density  and  is  therefore 
reduced  with  increased  altitude.  The  combined  effect  of  temper- 
ature and  density  changes  is  to  decrease  the  heat  transfer  but 
not  as  rapidly  as  the  engine  power  diminishes;  consequently 
the  cooling  capacity  of  the  radiator  becomes  excessive  at  high 
altitudes  and  may  be  more  than  double  the  required  capacity. 

As    the   head    resistance  

falls  off  more  rapidly  than 
the  heat  transfer  the 
"  figure  of  merit "  of  the 
radiator  increases  with 
altitude. 

From  the  above  con- 
clusions the  performance 
of  a  radiator  at  any  alti- 
tude can  be  calculated 
when  its  ground  perform- 
ance is  known.  For  ex- 
ample, take  the  flat  plate 
core  (E-8)  for  which 
ground  data  are  given  in 
Tables  17  and  18.  It  is 
desired  to  calculate  its  performance  in  summer  at  10,000  ft. 
altitude  and  120  miles  per  hour.  The  ground  data  are: 

Mass  flow  of  air  at  120  miles  per  hour  =  10.97  Ib.  per  square 
foot  per  second. 

Head  resistance  at  120  miles  per  hour  =  12.5  Ib.  per  square 
foot. 

Weight  of  core  and  contained  water  =  14.15  Ib.  per  square 
foot. 

The  mean  temperature  of  the  water  in  the  radiator  may  be 


V 

^ 

ov 

\ 

\ 

^ 

^ 

e-  4U 
ft. 

V 

^^ 

&, 

^ 

r^ 

\ 

<o     OJ 

o 

f     o 

^ 

*»  ^ 

x» 

\ 

1 

X 

X 

X 

X 

\ 

I    L° 
£ 

-40 

X 

\ 

tA 

^\ 

X 

GO 

3         A 

[          I 

5         I 

Z        l< 

;      z 

0        2 

4-       2 

3 

Altitude  in  Thousands  of  Feet 

FIG.  275. — Variation  of  air  temperatures  with 
altitude. 


366 


THE  AIRPLANE  ENGINE 


assumed  to  be  30°F.  below  the  boiling  point.  The  pressure  at 
10,000  ft.  is  10.2  Ib.  per  square  inch  (see  p.  389)  and  the  cor- 
responding boiling  point  is  194.2°F.  The  summer  mean  temper- 
ature at  10,000  ft.  may  be  taken  as  45°F.  (Fig.  275).  The 

mean  temperature  differ- 
ence at  10,000  ft.  will  be 
194.2  -  -  30  --  45  = 
119.2°F.  The  air  den- 
sity at  the  same  ele- 
vation is  0.0545  Ib.  per 
cubic  foot  (see  Fig.  276). 
The  mass  flow  at  10,000 

ft.  =  10.97  X 


04         &        IZ        16       ZO       24        2& 
Alfi-f-ude    in  Thousands  of  Feet 

FIG.    276. — Variation   of   air   densities   with 
altitude. 


7.98  Ib.  per  square  foot 
per  second.  The  energy 
dissipated  at  mass  flow 
of  7.98  Ib.  is  40  h.p.  per 

square  foot  per  100°  F.  temperature  difference  (see  Table  18); 

with  the  increased  temperature  difference  the  energy  dissipated 

119.2 

—  47.7    h.p.  per   square  foot. 


becomes  40  X  ~ 


100 


The  head 


resistance  (see  Table  18)  =  12.5  X  ^0750  =  9*°9  lb*  per  s^uare 
foot. 

The  degree  of  masking  required  at  altitudes  may  be  readily 
calculated  if  the  engine  h.p.  is  assumed  proportional  to  the  air 
density.  If  the  radiator  is 
just  adequate  in  level  flight 
at  a  given  speed  at  the 
ground,  it  will  be  capable  of 
more  cooling  than  is  required 
of  it  in  level  flight  at  the  same 
speed  at  higher  altitudes.  It 
is  therefore  possible  to  mask 
an  increasing  fraction  (and 
cut  down  thereby  the  mass  FIQ  277>_Radiatormaskingat  altitudes, 
flow)  as  altitude  increases. 

The  curve  of  Fig.  277  shows  how  much  masking  is  possible  for 
the  flat  plate  radiator  E-S  at  120  miles  per  hour,  but  the  curve 
is  practically  the  same  for  other  cores  and  speeds. 


THE  COOLING  SYSTEM 


367 


It  should  be  remembered  that  climbing  should  be  considered 
as  well  as  level  flight  in  any  discussion  of  radiators  and  of  mask- 
ing. The  speed  for  maximum  climb  may  be  only  one-half  that 
of  level  flight  at  certain  altitudes,  and  the  cooling  must  be 
adequate  for  the  climbing  condition.  This  consideration  alone 
would  require  a  masking  of  50  per  cent  for  such  planes  in  level 
flight.  If  the  relation  between  maximum  climbing  speed  and 
level  speed  is  known,  and  also  the  change  in  engine  revolutions 
and  power,  the  mass  flow  of  air  can  be  determined  under  both 
conditions  and  the  desirable  degree  of  masking  can  be  found. 


Valve 


Pump 


.Pump 


To  connect  systems 
in  series,  turn  A" 
through 30° 


Radiator 


Engine 


Radiator 


FIG.  278. — Radiator    interconnections    for    dual    engines    on    lighter-than-air 

machines. 

The  twin-engined  dirigible  offers  a  special  case  of  importance. 
Such  a  craft  may  operate  for  long  periods  with  one  engine  only, 
which  therefore  operates  at  low  speed  but  full  power.  If  the 
radiator  is  designed  for  maximum  speed  each  radiator  will  be 
too  small  for  its  engine  at  this  reduced  speed.  To  obviate  the 
use  of  a  larger  radiator  the  installation  may  be  arranged  as  in 
Fig.  278  in  case  the  water  pumps  are  of  such  construction  as  to 
permit  the  water  to  pass  through  when  they  are  idle.  Turning 
the  valve  A  through  90  deg.  will  circulate  the  water  through 
both  radiators  and  through  the  jackets  of  both  engines,  and  will 
thereby  prevent  the  idle  engine  from  freezing  up  and  will  give 
more  than  adequate  radiating  surface.  Some  provision  for 
masking  the  radiator  is  especially  desirable  in  this  case. 


368  THE  AIRPLANE  ENGINE 

Masking  can  be  partially  accomplished  by  varying  the  water 
flow,  as  by  by-passing  some  of  the  water  from  radiator  inlet  to 
outlet.  The  effect  of  reducing  the  quantity  of  water  is  to  reduce 
the  mean  temperature  of  the  water  and  thereby  to  reduce  the 
mean  temperature  difference  between  air  and  water.  The 
possible  range  of  control  by  this  means  is  small.  Shutters  across 
the  radiator  front  answer  the  purpose  more  fully,  although 
they  add  to  the  head  resistance.  They  may  be  operated  by  the 
pilot,  or  as  in  some  German  planes,  may  be  under  the  automatic 
control  of  an  electrical  resistance  thermometer.  Closed  shutters 
on  a  nose  radiator  decrease  the  head  resistance:  on  a  free  air 
radiator,  they  increase  it.  A  retractable  side  or  bottom  radia- 
tor, which  may  be  drawn  within  the  body  to  decrease  the  cooling 
effect,  is  occasionally  used.  It  may  be  arranged  most  conven- 
iently as  an  auxiliary  radiator  in  series  with  a  fixed  main  radiator 
which  has  no  masking  device  and  is  adequate  for  high-altitude 
evel  flight.  The  auxiliary  radiator  is  retracted  as  altitude 
is  gained.  The  increased  water  resistance  from  two  radiators 
in  series  is  objectionable.  Yawing  is  another  possibility. 

Effects  of  Yawing  Airplane  Radiators. — The  air  stream  does 
not  always  approach  the  radiator  at  right  angles  to  its  face. 
The  most  common  causes  of  this  are : 

1.  Radiator  mounted  in  the  propeller  slip  stream  where  the  air  strikes  the 
radiator  at  angles  other  than  normal  to  its  face. 

2.  Radiator  mounted  in  the  wing  (or  other  position)  where  the  axes  of  its 
passages  for  the  air  are  not  parallel  to  the  direction  of  motion  of  the  plane. 

3.  Radiator  pivoted  about  an  axis  perpendicular  to  the  direction  of  motion 
of  the  airplane  for  the  purpose  of  changing  its  inclination  for  the  regulation 
of  cooling  capacity  (masking). 

The  effects  of  yawing  a  radiator  through  angles  from  0  to 
45  deg.  are  (1)  to  decrease  slightly  the  mass  flow;  (2)  increase 
the  head  resistance  by  as  much  as  50  per  cent  in  the  case  of  cores 
of  low  head  resistance  but  much,  less  in  the  case  of  high-resistance 
cores;  and  (3)  in  some  cases,  for  angles  up  to  20  or  25  deg.  to 
increase  slightly  the  heat  transfer.  These  effects  vary  largely 
with  different  types. 

The  complete  radiator  consists  not  only  of  the  core  but  of 
top  and  bottom  headers.  The  top  header  may  serve  merely  as  a 
distributor  or  it  may  have  sufficient  capacity  to  serve  as  reserve 
and  expansion  tank  also.  The  latter  practice  reduces  complica- 
tions and  is  therefore  used  on  small  machines  intended  for  short 


THE  COOLING  SYSTEM 


369 


flights.  For  large  machines  used  for  long  flights,  an  adequate 
water  capacity  would  entail  a  large  frontal  surface  and  excessive 
head  resistance  of  the  header.  The  desirable  reserve  capacity 
in  British  practice  is  given  by  the  formula: 


Gallons 


h.p.  X  (endurance  in  hours) 
1,600 


f-fbfn 


Baffle    Plate 
Instrument 
Flange ^ 

«x*£3" 


9aOkfMt 


Supporting 
Brackets 


^"Shutter 
Bracket 

FIG.  279. — Details  of  typical  nose  radiator. 

The  reserve  water  tank  is  often  located  in  the  upper  wing  but 
there  is  danger  of  freezing  unless,  as  in  Fig.  274,  the  water  circula- 
tion is  through  the  tank;  the  objection  to  including  it  in  the 
circulation  is  the  increased  length  of  pipe  through  which  the 
water  has  to  be  forced. 

The  lower  header  is  a  collector  only  and  should  be  as  small  as 
practicable.  Both  headers  should  be  stream-lined.  The  headers 
and  their  contents  will  usually  add  50  per  cent  to  the  weight 
of  the  core  and  its  contents.  Occasionally  (as  in  the  Maybach 

24 


370  THE  AIRPLANE  ENGINE 

plant)  the  headers  are  divided  into  halves  by  vertical  baffles 
on  the  fore  and  aft  line.  Water  enters  the  left-hand  side  of 
the  lower  header,  passes  to  the  left-hand  side  of  the  upper  header, 
then  over  the  baffle  to  the  right-hand  side  and  down  to  its  exit  at 
the  right-hand  side  of  the  lower  baffle;  this  arrangement  causes 
greatly  increased  water  resistance  if  the  same  weight  of  water  is 
circulated;  if  the  weight  of  water  is  halved  so  as  to  maintain  the 
same  velocity  in  the  radiator  passage,  the  mean  temperature 
difference  between  air  and  water  will  be  diminished,  necessitating 
the  use  of  a  large  radiator. 

A  complete  nose  radiator  is  shown  in  Fig.  279.  Among  the 
details  to  be  noted  are  the  filler,  inlet  and  outlet  pipes;  the 
perforated  baffle  plate  between  the  inlet  and  the  upper  tank; 
the  overflow  pipe;  the  upper  and  lower  supporting  brackets;  and 
the  shutter  brackets.  The  filler  cap  is  of  hard  rubber  with  a 
safety  chain ;  a  better  construction,  avoiding  loss  from  the  snap- 
ping of  the  chain,  is  with  a  hinged  cap  held  closed  by  a  snap  wire. 

Pumps. — As  previously  pointed  out  (p.  363)  the  cooling  water 
required  is  not  more  than  J£  gal.  per  brake  horse  power  per 
minute.  Ordinarily  it  is  J4  gal.  per  minute  or  less.  The  re- 
sistance to  the  circulation  of  the  water  is  chiefly  in  the  radiator, 
but  is  considerable  in  other  parts  of  the  system;  its  magnitude  is 
variable,  but  may  be  assumed  to  be  from  4  to  8  Ib.  per  square 
inch  in  good  installations. 

The  water  horse  power  of  the  pump  of  a  100-h.p.  engine  using 
^  gal.  (2  Ib.)  of  water  per  horse  power  per  minute  against  8  Ib. 

2  X  100  1 

per   square    inch    pressure    is    8  X  144  X  — ™ —  X  QQ  QQQ  = 

0.116  h.p.  If  the  efficiency  of  the  pump  and  its  drive  is  20  per 
cent,  the  horse  power  used  to  drive  the  pump  will  be  0.166  -r- 
0.2  =  0.58  h.p.,  which  is  a  very  small  fraction  of  100  h.p. 
Consequently,  the  water  pump  efficiency  is  comparatively 
unimportant  and  the  type  selected  should  be  one  of  maximum 
simplicity  and  minimum  weight.  The  single  impeller  volute 
centrifugal  pump  meets  these  conditions  best  and  is  univer- 
sally used. 

In  a  volute  pump,  water  enters  axially,  is  caught  by  the  im- 
peller blades  and  is  given  a  high  velocity  of  rotation  before  it  is 
discharged  into  the  volute  casing,  from  which  it  escapes  through 
one  or  more  outlets.  The  number  of  outlets  is  usually  the  same 
as  the  number  of  banks  of  cylinders.  In  order  to  keep  down  the 


THE  COOLING  SYSTEM  371 

size  and  weight  of  the  pump  the  impeller  rotates  at  a  speed  greater 
than  that  of  the  engine;  one  and  one-half  engine  speed  is  common. 
If  the  impeller  blades  are  radial  (Fig.  280)  the  theoretical  dis- 
charge pressure  in  feet  of  water  is  given  by  V2/2g  where  V  is  the 
tip  speed  of  the  impeller.  Taking  an  impeller  diameter  of  4  in. 

/         4       2,400\2 
and  a  speed  of  2,400  r.p.m.  this  becomes  ( ir  X  TO  X     QQ  J    -f- 

2g  =  32  ft.  =  13.8  Ib.  per  square  inch.  The  water  velocity 
leaving  the  impeller  cannot  be  converted  completely  into  pressure 
head  and  there  are  various  impeller  and  casing  losses,  so  that  the 


Outlet 


Inlet 


FIG.  280. — Water  pump  of  Liberty  engine. 

actual  discharge  pressure  will  be  much  less  than  that  calculated 
above;  it  would  probably  be  less  than  one-half  the  theoretical 
value. 

The  resistance  to  be  overcome  by  the  pump  is  entirely  fric- 
tional,  and  varies  as  the  square  of  the  amount  of  water  circulated. 
The  amount  of  water  circulated  is  proportional  to  the  speed  of  the 
pump.  The  work  done  by  the  pump  is  proportional  to  the 
volume  of  water  circulated  multiplied  by  the  resistance,  or  is 
proportional  to  the  cube  of  the  pump  speed. 

The  pump  of  the  Liberty  engine,  Fig.  280,  has  a  2-in.  inlet 
and  two  outlets.  It  runs  at  1^  times  engine  speed,  and  has  a 
capacity  of  86  gal.  per  minute  at  2,000  r.p.m.  of  the  engine.  The 
impellers  are  radial  and  are  partly  shrouded.  The  packing  of  the 


372 


THE  AIRPLANE  ENGINE 


impeller  shaft  against  waiter  leakage  is  kept  compressed  by  a 
coiled  spring.  The  King-Bugatti  (410  h.p.)  pump  (Fig.  281) 
has  impellers  completely  shrouded  on  one  side.  As  there  is  only 
one  outlet  the  casing  is  of  complete  volute  form.  The  impeller 
is  5%  in.  diameter  with  eight  vanes,  the  web  being  drilled  to 
equalize  the  water  pressure.  The  shaft  is  packed  with  graphited 
asbestos  rope  packing,  held  under  compression  by  a  coiled  spring. 
The  inlet  is  2}^  in.  in  diameter;  the  outlet  is  2% 6  m-  It  is 
coupled  direct  to  the  engine  shaft. 


FIG.  281. — Water  pump  of  Bugatti  engine. 

The  Austro-Daimler  (200  h.p.)  pump  weighs  7.6  lb.,  has  an 
impeller  4.4.  in.  in  diameter,  inlet  and  outlet  diameters  1.42 
in.,  and  a  ratio  of  pump  to  engine  speed  of  1.89.  It  is  driven  from 
the  rear  end  of  the  crankshaft  by  a  bevel  £  ear  which  is  integral 
with  a  sleeve  forming  an  extension  shaft  (EX  282).  The  pump 
bevel  gear  floats  on  the  end  of  the  pump  spindle,  and  is  fitted 
with  a  large-diameter  thrust  ball-race  and  retaining  spring,  which, 
being  at  the  bottom  end  of  the  spindle,  are  as  far  away  as  possible 
from  the  impeller.  Both  the  pump  spindle  bearings  are  lubricated 
through  two  holes  drilled  in  the  pump  body  and  oil  grooves  cut 
in  the  spindle  bearings.  The  impeller  is  formed  with  six  vanes 


THE  COOLING  SYSTEM 


373 


and  is  completely  shrouded ;  it  is  keyed  to  the  spindle  and  secured 
by  a  gun-metal  nut  and  washer.     A  conically-faced  shoulder  is 


fn/ef. 


FIG.  282. — Water  pump  of  Austro-Daimler  engine. 

machined  on  the  pump  directly  beneath  the  impeller.     This 
shoulder  beds  into  the  bevelled  face  of  the  bronze  bearing,  form- 


50 


J- 

?  20 
1    ,0 


1000 


1200 


1400 


1800 


2000 


1600 

R.D.m 

FIG.  283. — Performance  curves  of  Austro-Daimler  water  pump.     > 

ing  a  water-tight  joint.     The  performance  curves  of  this  pump 
are  given  in  Fig.  283. 


374 


THE  AIRPLANE  ENGINE 


The  Maybach  (300  h.p.)  pump  (Fig.  284)  has  an  impeller 
4.46  in.  in  diameter,  inlet  2.13  in.  in  diameter,  outlet  1.97  in.  in 
diameter,  and  a  ratio  of  pump  to  engine  speed  of  2.  The  pump 

spindle  is  driven  through  a 
dog  clutch  at  its  lower  end 
by  a  short  vertical  spindle 
running  in  a  bronze  bush- 
ing; this  spindle  is  driven 
by  a  bevel  gear  meshing 
with  the  main  bevel  fixed 
on  the  rear  end  of  the 
crankshaft.  The  top  por- 
tion of  the  pump  spindle 
bearing  is  cupped  to  form 
the  housing  for  a  thrust 
ball-race,  above  which  is 
fixed  the  impeller.  The 
impeller  is  a  gun-metal 
casting,  having  six  helical 
vanes.  The  lower  half  of 
the  pump  body 


FIG.  284. — Water  pump  of  Maybach  engine. 


is    an 


aluminum  casting,  to  the  inlet  passage  of  which  the  diagonal 
water  pipe  from  the  radiator  is  coupled  by  a  rubber  con- 
nection. The  top  half  of  the  water  pump  body,  which  is  a  gun- 
metal  casting,  is  formed  with  six  helical  passages  leading  in 


10 


90       100      110      120       J30 


SO       60       70 
Gallons  per  Minute 

FIG.  285. — Performance  curves  of  Maybach  water  pump. 

a  reverse  helical  direction  to  the  impeller.  These  passages 
connect  with  the  common  vertical  outlet  passage  in  the  top  of  the 
body  casting.  The  center  portion  of  the  top  body  casting,  inside 


THE  COOLING  SYSTEM  375 

the  helical  passages  above  the  impeller,  is  domed  and  fitted  with 
a  screwed  plug.  This  plug  is  drilled  with  a  small  hole,  to  prevent 
an  air-lock.  Two  other  holes  are  also  drilled  in  the  bottom  of 
the  impeller  between  the  vanes  for  the  same  purpose.  The 
steel  ball  thrust  race  is  exposed  to  the  flow  of  water,  a  disadvan- 
tageous feature.  Performance  and  efficiency  curves  for  this 
pump  at  2800  r.p.m.  are  given  in  Fig.  285.  It  will  be  seen  that 
the  maximum  pump  efficiency  of  28  per  cent  is  obtained  with  a 
discharge  head  of  about  22  ft.  of  water  and  a  capacity  of  100  gal. 
per  minute. 

Piping. — Water  velocities  in  pipes  vary  from  about  8  ft.  per 
second  in  small  engines  to  16  ft.  per  second  in  large  engines. 
Actual  pipes  sizes  are  from  1}^  in.  diameter  for  90  h.p.  to  2  in. 
diameter  for  400  h.p. 

The  frictional  resistance  to  flow  of  water  through  straight 
pipes  is  given  by  h  =  4f  (l/d)  (V2/2g)  where  h  is  the  loss  of  head 
in  feet,  I  and  d  are  the  length  and  diameter  of  the  pipe  respec- 
tively in  feet,  V  the  velocity  in  feet  per  second,  and  /  is  a  coeffi- 
cient whose  value  is  likely  to  vary  from  0.004  to  0.010,  depending 
on  the  roughness  of  the  pipe.  Inlets,  outlets  and  bends  will  each 
offer  a  resistance  equivalent  to  a  length  of  10  to  20  diameters. 

Large  pipe  sizes  diminish  the  resistance  and  work  of  the 
pump  but  they  weigh  more  and  hold  more  water.  The  pump 
suction  should  be  of  ample  diameter  and  as  short  and  direct  as 
possible.  The  connecting  rubber  hose  should  be  firm  and  non- 
collapsible.  Pipe  lines  should  be  of  light  tubing,  bent  to  easy 
radii,  with  a  minimum  of  bends  and  fittings.  Hose  connections 
at  junction  points  should  be  very  short,  and  should  fit  over  cor- 
rugations. The  fastenings  should  be  by  smoothly-bearing  steel 
clamps  which  do  not  cut  the  rubber.  Tape  should  be  applied 
over  hose  and  clamp  and  the  whole  shellacked.  The  pipes  should 
be  arranged  to  avoid  air  pockets  if  possible;  if  such  occur,  vent 
cocks  must  be  applied.  Particular  care  must  be  given  to  the  vent 
cock  on  the  pump  casing. 

Water. — The  water  used  should  be  free  from  lime.  Filling  the 
system  with  boiling  water  makes  starting  easy  in  cold  weather. 
Anti-freezing  solutions  are  all  more  or  less  objectionable,  and 
it  is  best  to  drain  the  system  when  the  plane  is  not  in  use.  Fig. 
286  shows  the  properties  of  some  anti-freezing  mixtures.  Alcohol 
lowers  the  boiling  point  and  makes  close  control  of  temperature 
essential;  the  strength  decreases  and  the  freezing  point  is  elevated 


376 


THE  AIRPLANE  ENGINE 


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fe 

\ 

^ 

1" 

«  ?o 

^ 

\ 

\ 

S   dg 
& 

JC 

^ 

X.  % 

v 

\ 

^ 

\ 

ID 

X 

\ 

\ 

^ 

Jv 

t 

X, 

x 

s 

-10       -5          0          5          10          15        ZO         25         30 
Freezing  Poitrf,  Decj.Fahr. 

FIG.  286. — Properties  of  anti-freezing  mixtures. 


'Expansion    \ 

Tank          '"  Center  Section 


To  Expansion  Tank 

j 
Honey  tomb         /  from  Engine 


"Airlnfake 
Shufors 


WirestoShutferi     \    " 

Control  Lever  i        ^"- Drain  Cock 


Confrpf 


•in  Cock 


. 

Fron-1-  Vi  ew  of 


Drain  Cock  Radio  for 

Side      View 
FIG.  287.— Cooling  plant  of  S  E-5  airplane. 


THE  COOLING  SYSTEM 


377 


as  the  alcohol  evaporates.  Periodical  tests  by  hydrometer  are 
advisable.  Glycerine  does  not  evaporate,  but  impairs  circula- 
tion and  is  detrimental  to  rubber.  The  glycerine  should  be 
stirred  slowly  into  the  water. 

Typical  complete  cooling  systems  are  shown  in  Figs.  287  and 
288.  Figure  287  is  for  a  180  h.p.  Hispano-Suiza  engine  in  a 
SE-5  plane  with  two  tubular  nose  radiators  at  the  sides  of  the 
engine  shaft,  each  30  by  7K  by  S1^  in.,  with  a  total  frontal  area 
of  450  sq.  in.  and  a  radiating  surface  of  12,700  sq.  in.  The  water 
capacity  is  83^  Ib-  an(*  the  flow  rate  30  gal.  per  minute.  The 
system  is  provided  with  an  .expansion  tank  which  occupies  the 

Filler  Cap 


•To  Motor 


'••ToMoh>r 
Top  View  of  Rad/crfor 


FIG.  288. — Cooling  plant  of  Le  Pere  airplane. 

leading  section  of  the  middle  panel  of  the  upper  wing.  A  small 
portion  of  the  water  leaving  the  cylinders  passes  around  the 
intake  manifold  and  is  then  returned  to  the  pump;  the  rest  of  it 
goes  through  the  radiator.  The  radiator  is  masked  by  shutters. 
Figure  288  shows  a  360-h.p.  Liberty  engine  in  a  Le  Pere  two- 
seater  plane  with  a  wing  radiator  in  the  center  section  of  the 
middle  panel  of  the  upper  wing.  The  radiator  is  31  in.  long,  27 
in.  wide  and  7  in.  deep;  has  a  frontal  area  of  783  sq.  in.  and^a 
radiating  surface  of  35,520  sq.  in.  The  water  capacity  is  41.6  Ib. 
and  the  flow  80  gal.  per  minute  Free  water  area  61.6  sq.  in.,  free 
air  area  1,247  sq.  in.,  weight  127  Ib.  The  water  pumped  around 
the  manifold  goes  to  the  radiator  before  returning  to  the  pump. 


CHAPTER  XV 


GEARED  PROPELLER  DRIVES 

A  well  designed  airplane  engine  develops  its  maximum  power 
at  a  speed  (r.p.m.)  considerably  in  excess  of  the  most  efficient 
propeller  speed.  In  order  to  combine  maximum  power  develop- 
ment with  most  efficient  utilization  of  that  power  it  is  necessary 
to  resort  to  a  geared  drive. 

Geared  drives  have  been  employed  in  a  number  of  successful 

installations.  A  German  analy- 
sis of  these1  is  the  basis  for  the 
discussion  which  follows.  The 
simplest  type  is  a  single -reduc- 
tion with  spur  gears  as  in  Figs. 
289  and  290.  In  the  Renault 
engine  (Fig.  289),  the  gear  ratio 
is  two  to  one  and  consequently 
can  be  used  for  driving  both 
camshaft  and  propeller  shaft; 
in  the  Hispano-Suiza  engine 
(Fig.  290)  the  gear  ratio  is  four 
to  three.  Gears  of  this  type 
show  heavy  wear.  A  design  for 
a  single  reduction  with  internal 
gear  is  shown  in  Fig.  291;  the 
internal  gear  housing  is  attached 
to  the  crankcase  by  an  eccentric 
centering  flange  which  permits 
accurate  adjustment  of  the  gears.  This  type  permits  great 
simplicity  but  there  is  difficulty  in  arranging  satisfactory  bear- 
ings on  both  sides  of  the  gear  wheels. 

With  single-reduction  gears  the  propeller  shaft  cannot  be  in 
the  same  axial  line  with  the  crankshaft;  when  this  arrange- 
ment is  desired  double-reduction  gears  must  be  used.  There 
are  many  possible  arrangements;  both  pairs  of  wheels  may  be 
fitted  with  internal  or  external  gears  and  in  addition  any  one  of 
the  three  shafts  may  be  fixed  while  the  other  two  drive  and  are 
driven  respectively.  Some  of  these  arrangements  are  shown 
^UTZBACH:  Technische  Berichte,  Vol.  Ill,  Sec.  3. 

378 


FIG. 


--•  Crank 
Shaft 

20T.P4.57r(fnmm) 


289. — Renault  single-reduction- 
gear. 


GEARED  PROPELLER  DRIVES 


379 


FIG.  290. — Hispano-Suiza  single-reduction-gear. 


FIG.  291. — Single-reduction  internal  gear. 


Iran  k  Shaft  anci  Propeller  Turning  in  the  Same    Crank  Shaft  and  Propeller  Turning! 

_P_ir_ect^n I  Opposite  Directions 

In  termed '   ' 


rection 


7  re  bnarr  turning  in  I  he  Same 
Direct/on  as  Crank  Shaff 


Intermec/icrh  Shaft  Turning 
•'-  Opposite  Or  reef  ion 


A 

Irrkrmediote 
Shafting 
Fixed  in 
Housing 


B 

Intermediate 
Sfrafr 
Revo/vi'ng 
wH-h      ^ 
Propeller 


Intermediate 
Shaft 
Revolving 
with  Crank 
Shaft 


FIG.  292.— Possible  arrangements  of  double-reduction  gears. 


380 


THE  AIRPLANE  ENGINE 


schematically  in  Fig.  292.  In  the  top  row  the  intermediate 
shaft  is  fixed;  in  the  second  and  bottom  rows  it  revolves  forming 
the  so-called  planetary  gears.  The  shaded  gears  are  fixed  and  do 
not  revolve.  Some  of  these  arrangements  offer  considerable  diffi- 
culties for  actual  construction,  notably  in  the  matter  of  pro- 
viding suitable  bearings  on  both  sides  of  the  gears;  in  others 
the  space  occupied  may  be  great  and  the  revolutions  of  the 
intermediate  shaft  very  high. 

A  simpler  arrangement  is  one  in  which  both  pairs  of  gears 
have  one  gear  in  common.  Schematic  outlines  of  such  reductions 
are  shown  in  Figs.  293,  294  and  295.  Figure  293  is  developed 
from  Ai.  Fig.  292;  Fig.  294  from  A4;  and  Fig.  295  from  B2. 
The  Rolls-Royce  planetary  gear,  Fig.  296,  is  an  actual  construc- 
tion of  Fig.  295.  The  three  revolving  intermediate  shafts  are 


JT 


FIG.  293.  FIG.  294.  FIG.  295. 

Double-reduction  gears  with  a  common  gear. 

carried  in  a  spider,  C.  The  internal  gear,  a,  on  the  crankshaft 
drives  the  three  gears,  6,  on  the  intermediate  shafts,  and  the 
three  gears,  c,  on  the  same  shaft  mesh  with  the  gear,  d,  which  is 
held  against  revolving  in  the  housing.  The  spider,  C,  revolves 
and  carries  the  propeller  shaft. 

The  advantage  of  the  double-reduction  gear  over  the  much 
simpler  single-reduction  gear  lies  in  the  perfectly  axial  trans- 
mission of  the  power,  from  which  the  best  condition  of  loading  of 
the  housing  (pure  torsion)  is  obtained.  When  the  power  is 
transmitted  through  two,  three  or  four  intermediate  gears  at  equal 
angles,  springing  of  the  gear  shafts  from  unequal  peripheral  forces 
or  inaccurate  tooth  forms  is  avoided.  Certain  arrangements  also 
make  it  possible  to  use  heavy  revolving  masses  (for  instance, 
those  of  the  intermediate  shafts  or  the  larger  internal  gears), 
thereby  improving  the  uniformity  of  transmission  and  avoiding 
reversals  of  tooth  pressure.  The  principal  advantage,  however, 
consists  in  the  fact  that  on  account  of  the  load  being  divided 


GEARED  PROPELLER  DRIVES 


381 


o 


382 


THE  AIRPLANE  ENGINE 


between  two  to  four  intermediate  gears  the  tooth  pressures  per 
unit  of  tooth  face  are  low.  Consequently  small  pitches  and  small 
gears  can  be  used  which  in  turn  have  smaller  construction 
defects  since  the  defects  of  manufacture  resulting  from  the  use 
of  inaccurate  dividing  wheels  increase  with  increasing  radius. 
The  disadvantage  of  the  double  reduction  gear  is  its  weight  and 
cost  and  the  need  for  exact  adjustment  of  the  intermediate 
shafts  if  all  the  gears  are  to  work  equally.  Furthermore,  a 
complicated  construction  is  necessary  to  ensure  a  solid  and 
secure  assembly  of  the  gear. 

To  obtain  and  keep  proper  adjustment  of  the  reduction  gearing 
as  wear  occurs,  it  is  necessary  to  fit  a  joint  between  the  crank  case 
and  the  gear  case,  or,  in  the  transmission,  between  crankshaft  and 


FIG.  297. — Rolls-Royce  single-reduction  gear. 

gear,  which  will  adjust  itself  automatically  while  running  or 
can  be  adjusted  in  assembly.  In  the  Rolls-Royce  planetary 
gear,  Fig.  296,  a  sliding  cross  linkage  is  used  in  a  fixed  housing. 
The  link,  (B)  and  e,  lies  between  the  outer  engine  housing  and  the 
intermediate  gear  wheel,  d,  which  is  held  in  the  housing.  Con- 
sequently the  gear  wheel,  d,  can  adjust  itself  and  always  remains 
concentric  with  the  crankshaft.  The  whole  set  of  planetary 
gears  also  remains  concentric  with  the  crankshaft — which  may 
shift  in  the  casing — but  not  with  the  casing.  The  forward 
bearing,  g  and  h,  must  be  adjusted  on  each  overhaul  of  the  engine 
by  the  screws,  /. 

In  the  Rolls-Royce  spur  gear,  Fig.  297,  the  upper  gear  can 
be  adjusted  by  eccentrically-set  ball-bearing  cages,  c  and  d,  and 
the  lower  gear  can  be  adjusted  on  the  engine  shaft  by  screws.  A 
universal  joint  is  used  between  the  crankshaft  and  the  gear,  a. 


GEARED  PROPELLER  DRIVES  383 

Many  difficulties  have  been  encountered  in  the  actual  oper- 
ation of  reduction  gears — principally,  fracture,  wear  and  heating 
of  the  gears. 

The  bending  stress  in  a  gear  tooth  of  the  common  involute 
form  may  be  taken  as 

W 
f  —  14  X  r-  approximately, 

where  W  is  the  load  in  pounds  on  the  gear  tooth,  b  is  its  width 
and  p  is  the  circular  pitch  in  inches.  The  mean  value  of  the 
loading  on  the  tooth  can  be  determined  from  the  known  engine 
power,  P,  and  the  speed,  V,  of  the  pitch  circle 

_  550  XP 
V 

The  maximum  loading  on  the  teeth  may  be  considerably 
greater  than  the  mean  loading  either  because  of  acceleration 
pressures,  resulting  from  incorrect  pitch  or  form  of  teeth,  or 
because  of  irregular  delivery  of  power  from  the  engine,  or  on 
account  of  reinforced  vibration  near  a  resonance  period  of  the 
shaft.  The  values  of  /  calculated  for  a  number  of  successful 
engines  run  from  30,000  to  about  40,000  Ib.  per  square  inch; 
the  material  used  is  generally  case-hardened  chrome-nickel  steel. 
These  high  stresses  are  calculated  on  the  assumption  that  all 
the  load  is  carried  on  one  tooth.  With  accurate  pitching  the 
deformations  of  the  loaded  tooth  will  transfer  load  to  the  next 
tooth.  With  oblique  teeth,  such  as  herring-bone  gears,  the 
tooth  pressure  is  distributed  on  an  oblique  line  running  from  the 
root  to  the  tip  and  the  bending  stress  is  thereby  reduced.  The 
stresses  are  worse  if  the  teeth  bear  unevenly  as  a  result  of  warping 
in  hardening,  untrue  keying  or  poor  forming. 

The  surface  pressure  of  the  opposing  curved  tooth  faces  must 
not  be  sufficient  to  squeeze  out  the  oil  film.  The  relative  sliding 
speed  of  straight-toothed  gears  is  zero  at  the  pitch  circle,  and 
the  oil  is  more  easily  squeezed  out  under  this  condition  than  when 
there  is  relative  motion.  The  bearing  pressure  is  given  by  the 

W 

expression  -JT-J,  where  d  is  the  diameter  of  the  relative  curvature  of 

the  teeth  at  the  rolling  circle.     With  involute  teeth  having  radii 
of  curvature  of  e\  and  62  at  the  rolling  circle, 

2       1         1 

-;  =  —   H •) 

d       a  ~  e2 


384  THE  AIRPLANE  ENGINE 

the  +  sign  applying  to  external  gears,  the  —  sign  to  internal 
gears.  Calculations  from  successful  engines  indicate  that  with 
hardened  gears  the  bearing  pressure  may  go  up  to  1,400  Ib.  per 
square  inch;  if  the  gears  are  not  hardened  it  should  not  exceed 
450  Ib.  per  square  inch.  With  internal  gears  the  bearing  pres- 
sures become  low  and  hardening  is,  as  a  rule,  unnecessary. 
Experience  with  roller  bearings,  where  hardened  rolls  run  between 
hardened  rings,  indicates  a  permissible  bearing  pressure  of  2,800 
Ib.  per  square  inch  or  more  at  low  peripheral  speeds;  if  the  rolls 
bear  directly  on  the  unhardened  shaft  the  value  falls  to  from 
150  to  300  Ib.  per  square  inch. 

With  oblique  toothed  gears  the  contact  shifts  with  great 
speed  from  side  to  side,  as  a  result  of  which  there  is  less  tendency 
to  squeeze  out  the  lubricating  film. 

Heating  of  the  gears  results  from  the  sliding  contact  at  the 
teeth  and  may  be  of  such  magnitude  as  to  lead  to  trouble.  The 
heat  is  best  carried  off  by  thermal  conduction  from  the  gears 
to  the  outer  casing,  but  if  this  is  not  sufficient  it  must  be  assisted 
by  oil  cooling.  The  lubrication  should  not  be  so  heavy  that  the 
oil  heats  up  through  churning;  this  may  occur  through  the  use  of 
wide  gears  which  catch  the  oil  and  force  it  out  sideways  with 
great  force,  or  through  locating  the  gears  very  close  to  the 
housing. 

For  smooth  running  it  is  necessary  that  there  should  be  no 
reversals  of  pressure  in  the  gears.  Four-cylinder  engines  give 
such  reversals  of  pressure  and  so  do  six-cylinder  engines  at  a  low 
torque,  or,  with  very  heavy  reciprocating  parts,  at  high  speeds. 
With  a  larger  number  of  cylinders  with  crank  angles  equally 
spaced  reversals  will  not  occur. 

Central  power  plants  have  been  used  on  several  planes.  The 
principal  advantage  which  they  offer  is  the  possibility  of  con- 
centrating power  plants  in  a  central  engine  room  (where  they  can 
be  under  constant  supervision)  and  the  resulting  reduction  of 
drag  of  the  complete  machine.  There  is  also  the  possibility 
of  reducing  the  number  of  mechanics  required  in  a  multi-engined 
plane.  The  disadvantages  are  the  loss  of  power  (possibly  5  per 
cent)  in  the  transmission  shaft  and  gears,  and  the  increase  in 
weight. 

Chain-driven  propellers  were  used  successfully  by  the  Wright 
Brothers  in  1903  and  by  others  later.  In  recent  years  the  chain 
drive  has  not  been  used  but  shaft  and  bevel  gears  have  been 


GEARED  PROPELLER  DRIVES  385 

employed  with  some  degree  of  success.  Siemens-Shuckert 
multi-engine  planes  of  several  sizes  have  used  bevel  gear  drives. 
The  largest  of  these  with  six  engines  and  four  propellers,  is  ar- 
ranged with  the  four  rear  engines  driving  the  two  rear  propellers 
at  half  engine  speed  and  the  two  front  engines  driving  the  two 
front  propellers  with  a  reduction  ratio  of  14  to  9.  The  couplings 
between  the  engines  on  the  main  transmission  gear  are  a  combina- 
tion of  friction  and  independent  couplings.  The  latter  enable  the 
engine  to  be  disengaged  and  stopped  if  damaged.  The  articu- 
lated transmission  shafts  are  connected  at  both  ends  through 
laminated  spring  couplings. 

The  main  difficulty  in  the  operation  of  shaft  drives  has  been  in 
the  setting  up  of  "torsional  resonance/'  which  has  caused  break- 
age of  shafts  and  universal  joints.  This  has  been  overcome 
by  the  use  of  a  flywheel  on  the  engine  and  a  special  clutch  which 
combines  a  dog  clutch  and  a  friction  clutch.  The  shaft  should 
rotate  at  engine  speed  and  the  gear  reduction  should  be  near  the 
propeller.  Some  trouble  has  resulted  from  " whirling"  of  long 
shafts  but  probably  because  bearings  have  been  placed  at  nodal 
points;  this  can  be  avoided.  With  these  difficulties  overcome 
it  is  doubtful  whether  the  expense,  weight  and  complication  of 
the  flywheels,  clutches,  shafts  and  gears  will  not  more  than 
counterbalance  the  advantages  of  a  central  engine  room. 


CHAPTER  XVI 
SUPERCHARGING 

Change  of  Engine  Power  with  Altitude. — The  indicated  work 
in  the  cylinder  of  a  gasoline  engine  is  the  product  of  the  heat  of 
combustion  of  the  fuel  by  the  thermodynamic  efficiency  of  the 
engine.  The  thermodynamic  efficiency  is  unaffected  by  the  air 
density  and  depends  only  on  the  ratio  of  compression.  The  heat 
of  combustion  is  determined  by  the  weight  of  fuel  which  can  be 
burned  and  this  depends  on  the  weight  of  air  admitted  and  con- 
sequently on  the  density  of  the  air.  All  other  conditions  remain- 
ing constant,  the  indicated  power  of  an  engine  would  vary  directly 
as  the  density  of  the  air. 

The  brake  horse  power  of  the  engine  is  the  difference  between 
its  indicated  power  and  the  power  required  to  overcome  engine 
friction.  At 'constant  engine  speed  the  friction  will  not  change 
greatly  with  the  air  density;  it  increases  with  lowered  temperature 
of  the  lubricant  and  decreases  with  lowered  pressures  at  rubbing 
surfaces.  If  the  frictional  resistance  remained  constant  the 
brake  horse  power  would  fall  off  much  more  rapidly  than  the 
indicated  power  at  high  altitudes.  For  example,  an  engine 
developing  100  i.h.p.  at  the  ground  will  give  85  b.h.p.  with 
15  friction  h.p.  Operating  in  air  at  one-half  ground  density 
the  theoretical  indicated  power  is  50  h.p.  and  with  15-h.p.  friction 
there  would  be  35  b.h.p.  The  brake  power  would  be  diminished 
in  the  ratio  35/85  =  0.412,  while  the  indicated  power  is  halved. 

The  actual  diminution  in  brake  power  is  not  as  great  as  the 
preceding  calculation  would  indicate;  the  conditions  are  complex 
and  not  susceptible  of  exact  calculation. 

The  friction  horse  power  is  partly  rubbing  friction  and  water 
and  oil  pump  work  and  partly  work  done  in  overcoming  throttling 
losses  at  the  intake  and  exhaust  of  the  gases.  The  former  losses 
may  be  assumed  constant  with  varying  air  density;  the  latter  may 
be  assumed  to  vary  directly  as  the  air  density.  If  the  total 
fricton  loss  is  15  per  cent  of  the  full  indicated  power  at  the 
ground,  and  the  throttling  losses  are  assumed  to  be  one-third  of 
the  total  friction  loss,  and  if  the  indicated  horse  power  is  propor- 

386 


SUPERCHARGING 


387 


tional  to  the  relative  air  density,  d,  then  the  brake  horse  power,  B, 
at  any  air  density  is  given  by 

„      d  -  0.10  -  0.05d  0.95d  -  0.10 


0.85 


0.85 


where  B0  is  the  brake  horse  power  at  the  ground.  The  following 
table  gives  horse  powers  calculated  for  different  altitudes.  It 
will  be  seen  that  the  brake  horse  power  is  nearly  proportional  to 


Altitude,  feet 

Relative 
air  density,  d 

Relative 
air  pressure 

Relative 
b.h.p.,  B/B. 

0 

1.0 

1.0 

1.0 

6,000 

0.829 

0.801 

0.808 

12,000 

0.694 

0.645 

0.660 

18,000 

0.581 

0.518 

0.532 

24,000 

0.485 

0.414 

0.424 

30,000 

0.411 

0.333 

0.342 

the  air  pressure.  Actual  tests  support  these  calculated  quantities 
for  altitudes  up  to  10,000  ft.;  for  high  altitudes  the  brake  horse 
power  does  not  decrease  as  rapidly  as  the  air  pressure  nor  so 
slowly  as  the  air  density  but  according  to  some  intermediate  law. 

Tests  made  at  the  Bureau  of  Standards1  show  a  variation  of 
brake  horse  power  with  barometric  pressure  as  in  column  7  of 
Table  20.  The  ratio  of  brake  power  to  air  density  is  given  in  the 
eighth  column.  It  is  seen  that  the  brake  power  falls  off  more 
rapidly  than  the  air  density  and  that  at  one-half  ground  density 
the  brake  power  is  about  0.43  time  the  ground  power. 

The  variations  with  air  density  of  the  mechanical,  volumetric 
and  thermal  efficiencies  of  the  Liberty  12  engine  and  the  Hispano- 
Suiza  300  at  a  speed  of  1,600  r.p.m.  are  given  in  Fig.  298. 

The  relative  horse  powers  of  Table  20  are  based  on  constant 
engine  speed.  They  may  more  properly  be  regarded  as  relative 
engine  torques.  Engine  speed  falls  off  with  increasing  altitude 
so  that  the  actual  horse  power  developed  falls  off  more  rapidly 
than  is  indicated  in  Table  20.  With  constant  revolutions  per 
minute  the  resisting  torque  at  the  propeller  diminishes  in  direct 
proportion  to  the  air  density  and  consequently  falls  off  less 

I4th  Annual  Report,  National  Advisory  Committee  for  Aeronautics,  1918, 
p.  502,  Fig.  6. 


388 


THE  AIRPLANE  ENGINE 


rapidly  than  the  engine  torque.  Since  the  engine  torque  is 
practically  independent  of  the  revolutions  per  minute  the  engine 
speed  will  diminish  as  altitude  is  gained  until  that  speed  is  reached 
at  which  propeller  torque  equals  engine  torque. 

The  actual  engine  power  at  any  altitude  is  given  by 


P  =  PG  X  K  X 


No 


where  PG  is  power  developed  at  the  ground, 
No  is  revolutions  per  minute  at  the  ground, 
N  is  revolutions  per  minute  at  altitude, 
K  is  the  quantity  in  the  seventh  column  of  Table  20. 

Barometric  Pressure  in  Cm. of  Ha.  Approximate 


74.4      63.7       523      43.6     36.0     29.9 


74.51       63.29       52.51    43.43    35.81  29.74 


0.030      0.070      0.060       0.050       0.040       0.030       O.OoO       0.070      0.060      0.050       0.040      0.030 

Air  Densi-hj  in  Lb.per  Cu.  Ff. 

Liberty  12.  Hispano-Suiza  300. 

FIG.  298.  —  Variation  of  engine  efficiencies  with  air  density. 

Table  20  shows  that  the  engine  power  at  constant  speed  is  almost 
exactly  proportional  to  the  barometric  pressure.  On  this  basis 
the  engine  power  at  an  elevation  where  the  barometer  is  B  cm.  is 


Supercharging.  —  The  diminution  in  power  of  a  gasoline  engine 
with  increasing  altitude  results  in  a  moderate  reduction  of  speed 
in  horizontal  flight.  If  greater  power  were  available  the  ground 
speed  could  be  maintained  at  all  elevations  or  exceeded,  if  desired. 
Much  effort  has  been  expended  in  attempts  to  prevent  or  reduce 


SUPERCHARGING 


389 


•5 


PQ 


fc 

O 

§ 

8 

a 


11 

O     V 

^3  T3  ' 

a>    t, 

II 


§3  J  J 

H! 
IP 


U3    iO   U3   »O 

'-i<N<N<N<NC>><N<N<NlN'-iOa>OOt^ 
OOOOOOOOOOO   O  O   05  O  Oi 


OOOOCiOiCiOiaiOiOOOOOOOOOO 
r-I    1H    rH    O    O    O    O    O    O    0    O    O    O    O    O 


t-o 


^OOC^OOe«5O5>Oi-HOO^J<'-lOO»OCO^H 

i-Idddddddo'ddddo'dd 


I       I       I       I 


O5    »0    (N    O5    CD    CO    «-<    OS    t      "5 


<N          >O   d          U5  1C          IO  id 

OitN-*OCOC^OO500COlC^COC^»^OOi 
CIC^C^OlC^Wi-li-li-lrHi-tT-lr-tT-ll-l 


i— <r-<CS|COCO^^lOCOCOt>-t^-OOO5 


390  THE  AIRPLANE  ENGINE 

this  falling  off  in  engine  power.  Such  falling  off  would  be  entirely 
avoided  if  steam  power  could  be  substituted  for  gasoline  power, 
since  the  boiler  and  condenser  pressures  would  be  independent 
of  the  barometer  pressure.  Attempts  to  design  a  light-weight 
steam  plant  have  not  been  successful;  there  is  no  difficulty  with 
engine  or  condenser  (which  takes  the  place  of  radiator),  but  it 
has  not  been  found  practicable  to  design  a  boiler  to  withstand 
high  steam  pressures  and  of  sufficiently  extended  heating  surface 
without  arriving  at  weights  which  are  prohibitive  for  airplane  use. 
Furthermore,  the  lower  fuel  economy  of  a  steam  plant  would 
necessitate  the  carrying  of  a  greater  weight  of  fuel. 

Two  general  methods  present  themselves  for  increasing  gas 
engine  power  at  high  altitude: 

1.  To  select  an  engine  so  large  that  it  will  give  the  desired  power  when 
running  with  wide-open  throttle  at  the  high  altitude  at  which  the  airplane  is 
intended  to  fly,  and  to  operate  it  at  partial  throttle  at  all  lower  altitudes. 

2.  To  select  an  engine  which  gives  the  desired  power  at  the  ground  and 
add  some  device  for  supplying  the  cylinder  with  air  at  a  pressure  greater  than 
the  barometric  pressure  when  desired.      This  process  is  known  as  pre- 
compression  or  supercharging. 

As  an  illustration,  suppose  it  is  desired  to  fly  at  20,000  ft. 
developing  400  h.p.  This  can  be  accomplished  either  by  install- 
ing an  engine  which  would  develop  800  h.p.  with  wide-open 
throttle  at  the  ground;  or  by  installing  a  400-h.p.  engine  provided 
with  a  supercharging  device  which  is  able  to  maintain  that  horse 
power  at  all  altitudes  up  to  20,000  ft.  If  the  large  engine  is  used 
the  engine  weight  will  be  increased.  An  estimate  made  of  the 
increase  of  weight  which  would  result  from  doubling  the  power  of 
a  Liberty  motor  by  doubling  the  piston  area  per  cylinder,  while 
keeping  the  stroke  constant,  indicates  this  increase  would  be  about 
40  per  cent.  If  the  power  were  doubled  by  doubling  the  number 
of  cylinders  the  weight  would  be  nearly  doubled.  It  should  be 
noted  that  if  the  engine  is  not  permitted  to  develop  more  than 
400  h.p.  at  any  elevation,  the  radiator,  water  pump  and  general 
cooling  system  will  not  be  larger  than  for  a  400-h.p.  engine.  If 
the  smaller  engine  is  used  the  engine  weight  will  also  be  increased 
by  the  addition  of  the  supercharging  apparatus  and  the  engine 
becomes  more  complicated. 

Oversized  Engine. — In  this  system,  the  greater  weight  of 
the  engine  is  offset  not  only  by  greater  simplicity  (as  compared 
with  a  supercharging  engine)  but  also  by  greater  economy.  Such 


SUPERCHARGING 


391 


engines  should  be  provided  with  an  automatically  controlled 
throttle  valve,  actuated  by  some  device  (generally  similar  to  an 
aneroid  barometer)  which  responds  to  changes  in  atmospheric 
pressure.  An  example  of  such  a  device  is  given  in  Fig.  299  in 
which  an  airtight  flexible  chamber  filled  with  air  at  low  pressure 
actuates  a  balanced  double-seated  throttle  valve.  If  the  throttle 
is  placed  before  the  carburetor,  the  top  of  the  float  chamber 
must  be  kept  in  communication  with  the  low-pressure  side  of  the 
throttle.  The  control  may  be  so  adjusted  as  to  give  constant 
horse  power  at  all  altitudes  up  to  that  at  which  the  throttle  is 
wide  open;  the  power  cannot  be  maintained  beyond  that  point. 
With  an  engine  so  operated  it  is  possible  to  use  a  higher  ratio 
of  compression,  without  danger  of  preignition,  than  with  an 
engine  which  has  wide-open  throttle  at  the  ground. 


Air  Tight 
Chamber 


_T  Outside  Afr  u 

TJ   and  Tnlef  from  Compressor 
FIG.  299. — Automatic  throttle  control  for  oversized  engine. 

With  constant  power  output  the  weight  of  the  charge  admitted 
per  cycle  will  be  approximately  constant  and  the  pressure  in  the 
cylinder  at  the  beginning  of  compression  is  also  approximately 
constant.  The  latter  quantity  is  actually  a  little  more  at  the 
ground  than  at  higher  altitudes  because  the  engine  is  exhausting 
against  a  higher  barometric  pressure  and  consequently  there  is  a 
greater  weight  and  pressure  of  burned  gases  remaining  in  the 
cylinder  to  be  mixed  with  the  new  incoming  charge.  Further- 
more,' the  efficiency  at  the  ground  would  be  lower  than  at  high 
levels  on  account  of  the  higher  back  pressure.  With  constant 
power  output  the  pressure  in  the  cylinder  at  the  beginning  of  com- 
pression would  be  less  than  7  Ib.  per  square  inch  at  20,000  ft. 
elevation,  and  probably  less  than  8  Ib.  per  square  inch  at  the 
ground.  That  is,  the  maximum  pressure  to  be  expected  at  the 
beginning  of  compression  is  8  Ib.  per  square  inch  as  compared  with 
14  Ib.  in  the  usual  engine.  This  results  in  lower  compression  and 


392 


AIRPLANE  ENGINE 


explosion  pressures.  Furthermore,  the  cylinder  temperatures 
are  lower  throughout  the  cycle  mainly  in  consequence  of  the 
smaller  amount  of  heat  developed  by  the  explosion  and  the 
greater  cooling  effect  of  the  water  jacket.  Under  these  conditions 
it  is  possible  to  employ  higher  compresssion  without  danger  of 


Meters 
3048  46Z7 


10,000  14,000  18000 

Altitude, Feet. 


22POO 


FIG.  300. — Effect  of  altitude  on  variation  of  engine  power  with  compression 

ratio. 

preignition.     Engines  have  been  operated  in  this  manner  with  a 
ratio  of  compression  as  high  as  7. 

The  employment  of  a  high  ratio  of  compression  will  increase 
the  available  power,  particularly  at  high  altitude.  This  is  shown 
clearly  in  Fig.  300,  which  gives  the  results  obtained  at  the  Bureau 
of  Standards  with  an  engine  supplied  with  three  different  sets  of 
pistons  to  give  different  ratios  of  compression.  The  curve  B  is  for 
a  compression  ratio  of  5.3,  which  is  here  regarded  as  standard. 
The  curves  A  and  C  are  for  compression  ratios  for  4.7  and  6.2 
respectively.  It  will  be  seen  that,  calling  the  horse  power  with 


Per  Cent 
no-ease  of  M.E.P 
3  in  5  u 

|V'/5W 

^'^ 

x^ 

r"*--'^ 

»V^%17 

X<J 

*'' 

i^-  — 

& 

§ 

=/JW 

^ 

^ 

V5                        5.5   5.65   5 
Compression  Rat 

8       6 

0 

1100    1200     1300     1400     15.00    1600 
R.P.M. 

FIG.  301. — Variation   of  engine   capacity   with   compression  ratio   and   engine 

speed. 

standard  compression  100,  at  all  altitudes,  it  is  increased  to  104^ 
at  20,000  ft.  with  the  high  compression  and  reduced  to  95  J4 
with  the  low  compression.  At  the  ground  the  corresponding  horse 
powers  are  102J4  and  96%.  German  tests  on  a  Benz  200-h.p. 
engine,  Fig.  301,  show  similar  increase  of  power  with  ratio  of  com- 


SUPERCHARGING 


393 


pression  and  show  also  that  the  actual  increase  may  be  greater 
than  the  theoretical  (air  cycle)  increase,  particularly  at  high 
engine  speed. 

The  increase  in  engine  size  necessary  to  maintain  ground 
power  is  inversely  as  the  density  of  the  air  at  the  altitude  up  to 
which  full  power  is  desired.  The  percentage  increase  is  shown  in 
Fig.  302. 

The  change  in  horse  power  actually  developed  with  varying 
compression  will  be  greater  than  these  constant-speed  tests 
indicate.  With  a  given  propeller  the  engine  speed  will  increase 


cw 

* 

^^ 

^^ 

^*^ 

^*^ 

^^ 

fM 

100 

^*^* 

^•^ 

^^-^« 

^-*^ 

.^-  — 

0                  4000                8000              12000               16000              ZOC 

AH-i+ude.  Feet 
FIG.  302. — Required  displacement  volume  of  oversized  engine. 

with  the  engine  torque  and  will  cause  a  further  increase  in  horse 
power.     French  tests  show  the  following  results: 


Hispano-Suiza 

LeRhone 

Clerget 

Ratio  of  compression  

5.3 

5.8 

5.18 

5.65 

6  58 

4  6 

5  2 

5  6 

Revolutions  per  minute.,  .... 

2,040 

2,070 

1,230 

1,260 

1,290 

1,290 

1,350 

1,360 

Brake  horse  power  

186 

195 

119 

126 

134 

123 

137 

144 

The  admission  of  inert  gases  with  the  explosive  mixture  is 
now  being  developed  as  a  means  of  maintaining  high  economy 
at  low  levels  with  an  oversized  engine.  The  inert  gas  is  cooled 
exhaust  gas.  Its  presence  will  permit  operation  with  full 
throttle  at  much  higher  compression  ratios  than  would  otherwise 
be  possible  and  consequently  with  higher  thermal  efficiency. 
As  elevation  is  gained  the  percentage  of  inert  gas  in  the  charge 
may  be  reduced  either  by  hand  control  or  automatically;  it  should 
be  possible  to  maintain  the  engine  power  at  high  altitudes  and 
to  operate  continuously  at  very  high  efficiency  by  this  device, 
(seep.^435). 


394 


THE  AIRPLANE  ENGINE 


Up  to  12,000  ft.  altitude  the  oversized  engine  (with  about 
46  per  cent  increase  in  volume)  is  preferable  both  in  respect  of 
total  weight  and  of  simplicity  of  construction  to  the  supercharged 
engine.  With  altitudes  in  excess  of  20,000  ft.  the  weight  of  the 
oversized  engine  becomes  considerable  and  the  preference  may 
rightly  fall  on  the  supercharged  engine.  A  combination  of  the 
two  may  possibly  turn  out  to  be  best,  with  power  maintained 
constant  up  to  about  10,000  ft.  by  the  gradual  opening  of  the 
throttle  valve,  and  then  bringing  into  action  a  supercharging 
device  to  maintain  constant  power  up  to  20,000  ft. 

Supercharging  Engine. — In  a  supercharging  engine,  air  is 
compressed  by  a  blower  or  other  device  and  is  delivered  to  the 


75       70       65       60       55       50      45      40       55       30       25 
Ex  Ho  us-!-  Back  Pressure  (Cms.ofHcj.) 

FIG.  303. — Chart  for  finding  the  power  developed  by  a  supercharged  engine. 

carburetor  at  a  pressure  in  excess  of  the  surrounding  atmospheric 
pressure  and  consequently  in  excess  of  the  exhaust  pressure, 
except  in  the  case  (discussed  later)  where  the  blower  is  driven  by 
an  exhaust  gas  turbine.  The  power  that  can  be  delivered  by 
an  engine  whose  admission  and  exhaust  pressures  are  different  is 
readily  calculable  for  an  ideal  engine,  but  it  is  necessary  to  have 
recourse  to  actual  tests  in  order  to  ascertain  its  magnitude  for  an 
actual  engine.  Such  tests  have  been  conducted  at  the  Bureau 
of  Standards;  the  curves  given  in  Fig.  303  show  the  results 
obtained.  These  curves  give  the  horse  power  that  will  be 
developed  by  an  engine  with  any  exhaust  pressure  from  76  to  20 
cm.  of  mercury  and  with  air  supplied  to  the  carburetor  at  any 
pressure  from  76  cm.  down  to  55  cm,  of  mercury.  The  horse 


SUPERCHARGING  395 

power  is  given  as  a  ratio  to  the  horse  power  delivered  at  the  ground 
with  admission  and  exhaust  both  at  a  pressure  of  76  cm.  of  mer- 
cury. The  tests  were  conducted  with  the  carburetor  adjusted  to 
give  maximum  power  and  the  curves  are  all  corrected  to  the  same 
air  temperature  at  the  carburetor.  The  curves  show  that  if  the 
pressure  at  the  carburetor  is  maintained  at  76  cm.  during  flight 
the  horse  power  developed  in  the  engine  will  increase  as  a  result 
of  diminishing  back  pressure  and  at  20,000  ft.  (36.5  cm.  pressure) 
will  be  about  6  per  cent  greater  than  the  horse  power  at  the 
ground;  this  is  to  be  compared  with  the  diminution  of  51  per 
cent  in  horse  power  (see  Table  20)  at  the  same  altitude  without 
supercharging. 

The  gain  in  engine  horse  power  with  supercharging  is  not  of 
course  net  gain.  Some  of  the  additional  work  is  used  up  in  pre- 
compressing  the  air  to  the  admission  pressure.  Furthermore,  the 
precompression  heats  up  the  air  so  that  it  enters  the  carburetor 
at  a  temperature  greater  than  that  of  the  surrounding  atmosphere. 
The  actual  horse  power  developed  in  the  cylinder  is  given  by  the 
equation 

PC  =  PG  X  r  X  F 

where  Pc  is  the  horse  power  developed  with  the  supercharging 
apparatus  at  the  given  altitudes;  PG  is  the  observed  horse  power 
on  the  ground  at  the  observed  carburetor  air  temperature,  i\]  r  is 
the  horse  power  ratio  at  the  given  condition  of  exhaust  and 
carburetor  pressures  produced  by  the  supercharging  device  at  the 
given  altitude,  (obtained  from  curves,  Fig.  303)  andF  is  the  temper- 
ature correction  factor  to  correct  from  observed  temperature 
at  the  ground,  ti,  to  temperature  at  the  carburetor,  t2. 

The  temperature  of  the  precompressed  air  can  be  calculated 
from  the  equation 

n-l 


n 


where  T2  is  the  absolute  temperature  of  the  air  entering  the 
carburetor,  T3  is  the  absolute  temperature  of  the  air  entering  the 
supercharging  device,  p2  and  ps  are  the  air  pressures  at  the  same 
places.  The  quantity  n  may  be  assumed  for  ordinary  conditions 

T 

to  have  the  value  1.3.     The  quantity  «-  may  be  obtained  from 

Fig.  304,  which  gives  values  also  for  n  —  1.2  and  n  =  1.4. 


396 


THE  AIRPLANE  ENGINE 


The  effect  of  this  increase  of  temperature  is  to  diminish  the 
power  of  the  engine.  The  temperature  correction  factor  is 
obtained  from  the  equation  (compare  p.  33). 

=  920  +  t,. 
920  +  *, 
where  t\  and  tz  are  in  degrees  Fahrenheit  (not  absolute). 

As  an  example  of  the  use  of  the  above,  suppose  an  engine 
develops  200  h.p.  at  the  ground  with  barometer  76  cm.  mercury 
and  temperature  40°F.  and  that  it  is  taken  to  an  altitude  where 
the  barometer  is  35  cm.  and  temperature  —  4°F.  and  that  it  is 
supercharged  to  65  cm.  pressure.  From  Fig.  303  the  horse  power 


''1.0    I.I     1.2    1.3    1.4    1.5    1.6    1.7    1.8    1.9    2.0  2.1    2.2    2.3   24  25 

jj 

FIG.  304. — Temperature  rise  of  air  during  compression. 


ratio,  r,  corresponding  to  these  conditions  is  0.9.     The  pressure 
ratio—  =  ™-?   =  1.78.     Assuming  thatn  =  1.3,  Fig.  304  gives 


™ 

OO.O 


TIT  =  1-142,  or  T2  =  1.142  X  (460  -  4)  =  520.     Consequently 

/  3 

£2  =  Tz  -  460  =  60°.     The  temperature  correction  factor  F  = 
920  4-  60  =  0-978-     The  engine  horse  power  will  then  be 
PC  =  200  X  0.9  X  0.978  =  176 

If  the  barometric  pressure  at  which  the  ground  horse  power  is 
observed  is  not  76  cm.,  a  further  correction  may  be  introduced. 
For  example,  if  the  barometer  reads  74  cm.,  the  horse  power  ratio 
as  compared  with  76  cm.  is  found  from  curve  E,  Fig.  303,  to  be 
0.972.  The  horse  power  actually  developed,  PC,  will  then  be 
176 
0.972  " 

As  previously  pointed  out,  this  horse  power  is  not  the  net 
horse  power  available  for  driving  the  propeller.  There  must  be 
subtracted  from  it  the  work  requiredjto  precompress  the  air. 


SUPERCHARGING 


397 


The  ideal  method  of  compression  is  isothermal  but  this  cannot  be 
realized.  If  there  were  no  addition  or  abstraction  of  heat 
during  the  compression  and  no  frictional  resistance  or  eddy  losses, 
the  compression  would  be  adiabatic  and  this  is  what  the  cen- 
trifugal compressor,  without  cooling,  might  be  expected  to  accom- 
plish. The  work  of  adiabatic  compression  is  given  by 


=  wR^Ti(T^-  r») 

=  wJ  Cp  (T,  -  T3) 

where  pz,  PB,  are  the  pressures  at  beginning  and  end  of  compres- 
sion respectively;  vz,  v$,  the  corresponding  volumes;  T2,  Tz,  the 


60        34 

s 

^ 

s 

/ 

MO.        1 

/^, 

_~2l 

X 

^,x 

x^*'x/ 

X 

vv^x"^ 

^x^. 

--• 

8  -S    80  -3 

*  V.  "o    74 

x'^x 

^t^ 

^"^ 

j* 

^x^ 

J*^^ 

7    |   TO  _§  H 

X^xX"! 

J-^x^ 

5£Sl^ 

x 

x^^-* 

x^c 

'l^-3 

» 

^X^x- 

•^x*^ 

llJlt 

X^x*x| 

.X^x* 

2        20  1§ 

x#^[^i 

0          10 

0°  1 

8000          12000         16000         20000       24000 
Altitude  in  Feet 


FIG.  305. — Power  absorbed  in  the  compression  of  air. 

corresponding  absolute  temperatures;  72  is  the  air  constant,  53.4; 
J  the  mechanical  equivalent,  778;  Cp  the  specific  heat  of  air  at 
constant  pressure,  0.241;  7  the  ratio  of  specific  heats,  1.4;  and 
w  the  weight  of  air  compressed.  The  final  temperature,  T3, 
is  given  by  the  equation 


/PA 

=  y 


p 

The  work  done  per  pound  of  air  compressed  (adiabatically)  per 
second  is  given  in  Fig.  305,  curve  Wy  for  compression  from  the 
mean  pressure  and  temperature  existing  at  any  altitude  to 
standard  atmospheric  pressure.  This  quantity  is  expressed  in 
curve  P  as  a  percentage  of  the  work  which  can  be  done  by  that 
air  in  an  efficient  engine.  The  ratio  of  compression  (pressure 
ratio)  is  shown  by  curve  C  and  the  temperature  rise  of  the  air  by 
T.  The  mean  temperatures  used  as  a  basis  for  calculating  the 
above  curves  are  given  by  Fig.  275. 


398  THE  AIRPLANE  ENGINE 

With  isothermal  compression  the  work  required  to  compress 
1  Ib.  of  air  is  given  by  W  =  RTS  loge  —  ft.-lb.  The  actual  work 

done  on  a  centrifugal  compressor  in  compressing  air  is  found  to  be 
about  twice  the  work  of  isothermal  compression.  For  the  exam- 
ple worked  out  above,  the  work  of  isothermal  compression  for 

1  Ib.  of  air  is  W  =   534   X  456   X  loge  1.78  =  14,040  ft.-lb. 
The  actual  work  of  air  compression  per  pound  of  air  will  be 

2  X  14,040  =  28,080  ft.-lb. 

The  amount  of  work  done  by  1  Ib.  of  air  in  the  cylinder  is 
determinable  from  the  assumption  that  the  explosive  mixture  is 
15  parts  air  to  1  part  gasoline,  by  weight,  and  that  the  fuel 
consumption  is  J-£  Ib.  gasoline  per  horse  power  hour.  Every 

pound  of  air  does  work  2  x  33,000_X_60  =  264)000  f t  _lb      Qf 

10 

28  080 
this  work  the  fraction  0^4  nflf)  =  0-1^65  is  used  for  precompres- 

sion.  Consequently  in  the  case  discussed  the  net  available 
horse  power  will  be  181  X  (1  -  0.1065)  =  161  h.p.,  that  is, 
20  h.p.  will  be  used  in  driving  the  blower  or  other  supercharging 
device. 

There  is  one  type  of  supercharging  device  in  which  the  power 
required  for  pre compressing  the  air  is  obtained  from  an  exhaust 
gas  turbine.  This  imposes  a  back  pressure  during  the  exhaust 
stroke  in  excess  of  the  atmospheric  pressure.  For  example, 
suppose  that  under  the  same  conditions  as  those  worked  out  in 
the  example  with  65  cm.  carburetor  pressure,  an  exhaust  gas 
turbine  is  used  and  that  the  exhaust  back  pressure  is  60  cm. 
mercury.  The  horse  power  ratio  is  now  (Fig.  303)  0.85  and  the 

o  8^ 

net  horse  power  is  -^-  X  181  =  171  h.p. 

In  the  exhaust  gas  turbo-superchargers  that  have  been  built 
up  to  the  present,  the  selected  operating  conditions  have  generally 
been  the  maintenance  of  ground  pressure  in  both  admission  and 
exhaust  manifolds  up  to  some  limiting  altitude.  Assume  a 
ground  pressure  of  76  cm.,  temperature  66°F.  and  the  engine 
operating  at  an  altitude  where  the  barometer  is  38  cm.  (19,000 
ft.),  and  air  temperature  5°F.  If  the  exponent  n  during  the 
compression  has  the  value  1.3,  it  is  seen  from  Fig.  304  that 

m 

Y^  =  1-174  for  2*  =  2.     As  T3  =  460  +  5,   T2  =  545,  or  the 


SUPERCHARGING 


399 


temperature  of  the  air  entering  the  carburetor  is  545  —  460  = 
85°F.     The  engine  horse  power  is  thereby  diminished  in  the 

S  -  °-984- 

The  work  W  available  from  the  exhaust  gas  turbine  may  be 
determined  from  the  equation  for  adiabatic  compression  given 
above.  The  velocity,  V,  with  which  the  gas  discharges  on  the 
blades  of  the  turbine  (assuming  a  frictionless  nozzle)  is  given  by 

V2 
the  equation  W  =  -~-,  or  it  may  be  obtained  from  the  equation 


V  = 


.  per  second. 


where  T2  is  the  absolute  temperature  and  p2  the  absolute  pressure  of 
the  exhaust  gases  entering  the  turbine  nozzle.  Values  of  V  from  this 
equation  are  given  in  following  table  calculated  for  T2  =  1,800. 
The  average  temperature  of  the  gas  leaving  the  engine  is  about 
1,500°F.,  but  loss  from  the  exhaust  manifold  reduces  it  to  about 
1,300°F.,  or  1,800°  absolute. 


5 

1.1 

1.2 

1.3 

1.4 

1.5 

1.6 

1.7 

1.8 

1.9 

2.0 

P3 

V 

754 

1,026 

1,230 

1,384 

1,514 

1,622 

1,716 

1,798 

1,871 

1,940 

With  an  exhaust  gas  turbine  of  100  per  cent  efficiency,  the  work 
that  can  be  done  by  the  gas  is  equal  to  the  kinetic  energy  of  the 
gas.  If  an  exhaust  gas  turbine,  maintaining  ground  pressure  at 
the  carburetor  and  in  the  exhaust  manifold,  is  fitted  to  a  200-h.p. 
engine,  using  0.5  Ib.  gasoline  per  horse  power  hour  and  15  Ib. 
of  air  per  pound  of  gasoline,  the  weight  of  exhaust  gases  will  be 

^p       6  =  26.7  Ib.  per  minute.     At  19,000  ft.  altitude 

(38  cm.  barometer)  the  pressure  drop  ratio  — -  in  the  expansion 
nozzle  is  2,  and  the  corresponding  gas  velocity  is  1,940  ft.  per 


second.     The  kinetic  energy  of  the  gas  is 

,       26,000 
ft.-lb.  per  second  or  =  47.3  h.p. 


26.7  X  1,9402 


=  26,000 


2g  X60 
Assuming  an  efficiency 

of  50  per  cent  for  the  gas  turbine  the  power  available  for  driving 
the  compressor  is  23.7  h.p.  The  compressor  work  with  isother- 
mal 'compression  would  be  W  =  534  X  465  X  loge  2  =  17,400 


400  THE  AIRPLANE  ENGINE 

ft.-lb.  per  pound  of  air  compressed.     The  weight  of  air  com- 

,       200  X  0.5  X  15 
pressed  is  -  An       A^  --  =  0.417  Ib.  per  second.     The  power 

OU  X  OU 

.     ,  ,      .     ,,  .  0.417  X  17,400 

required  for  isothermal  compression  is  -          --  -  =  13.2 


h.p.  If  the  compressor  efficiency  is  50  per  cent  as  compared 
with  isothermal  compression,  the  power  required  to  drive  the 
compressor  will  be  2  X  13.2  =  26.4  h.p. 

For  the  operation  of  the  turbo  compressor  just  discussed, 
it  is  necessary  that  the  over  -all  efficiency  of  the  combination 

13  2 

should  be  not  less  than  -7^5  =  0.279.     This  over-all  efficiency,  E, 
4/  .o  . 

is  the  product  of  the  turbine  efficiency  ET  and  the  compressor 
(isothermal)  efficiency  Ec,  or  E  =  ET  X  Ec.  Tests  on  the 
Rateau  exhaust  gas  turbo-compressor  indicate  the  possibility  of 
values  of  ET  of  about  0.53,  and  a  value  of  about  0.5  for  Ec. 
These  correspond  to  E  =  0.53  X  0.5  =  0.265.  The  theoretical 
work  of  isothermal  compression  which  can  be  done  by  this 
combination  for  the  special  case  under  discussion  is  47.3  X  0.265 
=  12.55  h.p.  or  is  less  than  the  13.2  h.p.  calculated  as  necessary 
to  maintain  ground  pressure  at  the  carburetor. 

The  actual  pressure  which  could  be  maintained  at  the  carbu- 
retor is  readily  calculable.  The  work  of  isothermal  compression 

per  pound  of  air  is  W  =  RT3  loge—  2;  and  as  the  weight  of  air  is 
0.417  Ib.  per  second,  the  horse  power  for  isothermal  compression 
is  (0.417  X  RT3  logc—  )/550  =  12.55.  The  value  of  T3  has  been 

7?9 

given  as  456.     Solving  the  equation  gives  --  =  1.95,  and  p%  = 

74  cm.  That  is,  the  power  developed  by  the  exhaust  gas  turbine 
is  sufficient  to  compress  the  air  to  74  cm.  pressure.  If  a  higher 
pressure  is  desired  at  the  carburetor  either  the  back  pressure 
on  the  engine  must  be  increased,  thereby  increasing  the  turbine 
power,  or  the  efficiencies  of  turbine  and  compressor  must  be 
increased. 

The  inefficiency  of  the  compressor  has  a  further  effect  on 
the  engine  performance  besides  that  just  discussed.  Practically 
all  the  work  done  on  the  compressor  goes  finally  into  heating  of 
the  air  and  as  the  (isothermal)  efficiency  of  the  compressor  is 
about  0.5  the  amount  of  such  heating  can  be  readily  determined. 


SUPERCHARGING  401 

For  the  case  under  discussion  with  air  at  38  cm.  pressure,  and 
465°  absolute  temperature  Fahrenheit  and  with  compression  to 
74  cm.  pressure,  the  work  of  isothermal  compression  per  pound 
of  air  is  53.4  X  465  X  loge  1.95  =  16,600  ft.-lb.  =  21.35  B.t.u. 
The  total  work  done  is  2  X  21.35  =  42.7  B.t.u.  and  the  conse- 

42  7        42 
quent  heating  of  the  air  is  -    -  =  r^pr  =  177°F.     The  final 

C 


temperature  of  the  air  will  be  177  +  465  =  642°  absolute  =  182°F. 

This  high  temperature  of  the  air  entering  the  carburetor  will 
cause  a  decrease  in  volumetric  efficiency  of  the  engine.  To  avoid 
the  consequent  loss  of  engine  power  the  air  should  be  partly  cooled 
on  its  way  from  the  compressor  to  the  engine.  Some  heating  of 
the  air  is  highly  advantageous  in  aiding  the  vaporization  of  the 
fuel  in  the  carburetor  and  intake  manifold. 

Centrifugal  compressors  (single  stage)  do  not  appear  to  be 
suitable  for  ratios  of  compression  greater  than  2  to  1  on  account 
of  the  excessive  speeds  which  become  necessary.  This  means 
that  they  can  be  used  only  up  to  altitudes  of  about  .20,000  ft.  if 
they  are  to  maintain  ground  pressure  at  the  carburetor.  Multi- 
stage compressors  permit  higher  ratios  of  compression,  or  the 
same  ratio  with  lower  peripheral  speeds. 

Supercharging  Devices.  —  Two  methods  have  been  employed 
for  supplying  the  engine  with  air  at  a  pressure  higher  than  that  of 
the  surrounding  atmosphere. 

1.  The   cylinder  takes  in   an  overrich  charge  in  the  usual 
manner  and  this  charge  is  raised  in  pressure  and  diluted  to  the 
proper  strength  by  the  admission  of  compressed  air  at  the  end  of 
the  admission  stroke. 

2.  The  whole  of  the  air  going  to  the  engine  is  compressed 
and  is  sent  under  pressure  through  the  carburetor. 

In  the  first  method  the  demand  for  compressed  air  is  intermit- 
tent and  the  compression  is  most  suitably  carried  out  in  a  recip- 
rocating (piston)  compressor.  The  second  method  requires  a 
centrifugal  compressor.  It  suffers  the  disadvantage  that  the 
pressure  in  the  carburetor  is  greater  than  the  external  pressure 
so  that  the  carburetor  must  either  be  made  strong  and  tight 
enough  to  withstand  this  condition  or  must  be  entirely  enclosed 
in  a  chamber  under  the  compressor  pressure,  which  makes  it 
comparatively  inaccessible. 

An  example  of  the  first  method  is  the  Ricardo  system,  which 
has  been  experimented  with  considerably  in  England.  The 

26 


402 


THE  AIRPLANE  ENGINE 


Hand  Opera  tea1 
Valve  Here 


In  far  Cooler 


Automatic  Air  Valve 


cylinder  (Fig.  306)  has  a  ring  of  ports  uncovered  near  the  lower 
end  of  the  stroke.  The  piston  is  of  two  diameters,  as  shown, 
leaving  an  annular  space,  A,  which  diminishes  in  volume  as  the 

piston  descends.  This  annular 
space  communicates  freely  with 
the  intercooler,  5,  which  con- 
nects with  the  ring  of  ports. 
The  closing  of  a  hand-operated 
valve  at  F  puts  the  super- 
charger out  of  action  when 
desired.  Air  is  admitted  to 
the  annular  chamber,  A, 
through  the  automatic  valve, 
E. 

In  Fig.  306  the  piston  is 
shown  near  the  end  of  the 
suction  stroke.  Compressed 
air  from  B  is  just  beginning  to 
flow  into  the  cylinder  through 
the  ring  of  ports;  the  inlet 
valve  is  nearly  closed  and  the 
pressure  in  the  cylinder  is 
raised.  During  the  succeeding 
compression  stroke  air  is  ad- 
mitted through  E  into  A  and 
is  compressed  in  A  and  B  dur- 
ing the  expansion  stroke.  Near 
the  end  of  the  expansion  stroke 
the  ring  of  ports  opens  and  some 
of  the  burned  gases  pass  into  B\ 
the  exhaust  valve  opens  im- 
mediately afterwards  and  these 
burned  gases  together  with  the  compressed  air  sweep  back 
through  the  ports  and  scavenge  the  cylinder.  A  new  charge  of 
air  is  taken  in  through  E  during  the  exhaust  stroke  and  is  com- 
pressed during  the  following  suction  stroke. 

It  is  obvious  that  this  system  can  be  used  only  for  moderate 
degrees  of  supercharging  since  the  additional  air  supplied  to  the 
cylinder  per  admission  cannot  be  greater  than 


FIG. 


306. — Ricardo    supercharging    en- 
gine. 


L  X  A(D*  *  d2) 


SUPERCHARGING 


403 


where  L  is  the  piston  stroke,   and  D  and  d  are  the  two  pis- 
ton diameters.     As  the  normal  cylinder  charge  has  a  volume 

L  X    D2  this   represents  an  increase   of  —     —  = 


(-=) 


If  d  =  J-£D,  the  increase  in  charge  would  be  75  per  cent.  The 
results  of  tests  with  this  device  are  given  in  Fig.  307;  they  show 
an  increase  of  power  of  approximately  50  per  cent. 

The  more  promising  method  of  supercharging  appears  to  be 
that  in  which  all  the  air  going  to  the  engine  is  precompressed  in 
a  compressor.  Reciprocating  compressors  operating  at  engine 
speed  have  been  tried  in  England  but  show  an  over-all  efficiency 
which  is  very  low  —  only  about  21  per  cent.  The  Roots  type  of 
positive  blower  (Fig.  308)  with  aluminum  rotors  operating  at 


FIG.  307. — Performance  curves  of 
^Ricardo  supercharging  engine. 


FIG.  308.— Roots  blower. 


twice  engine  speed  gives  an  over-all  efficiency  of  about  52  per 
cent  but  is  exceedingly  noisy.  The  most  commonly  used  type 
is  the  centrifugal  compressor.  Such  a  compressor  may  either  be 
directly  coupled  to  the  engine  or  driven  by  a  separate  engine 
or  by  an  exhaust  gas  turbine.  In  any  case,  its  peripheral  speed 
must  be  high  in  order  to  keep  down  the  number  of  stages  and  the 
weight  and  bulk  of  the  compressor.  If  directly  coupled  to 
the  engine  shaft  a  train  of  gears  must  be  employed  to  increase  the 
speed  up  to  10,000  r.p.m.  or  more,  depending  on  the  number  of 
stages  employed  and  the  amount  of  supercharging  desired;  if 
operated  through  a  gas  turbine  no  gears  are  necessary  as  the 
turbine  speed  will  be  from  20,000  to  30,000  r.p.m.  and  one 
compressor  stage  will  be  sufficient  at  these  speeds. 

Geared  direct-coupled  compressors  have  given  much  trouble 
from  stripping  of  gear  teeth.  At  the  high  speeds  of  rotation 
required,  the  kinetic  energy  of  the  rotor  wheels  is  very  great  and 


404 


THE  AIRPLANE  ENGINE 


considerable  forces  have  to  be  employed  for  rapid  acceleration. 
When  the  engine  is  started  or  the  throttle  valve  is  opened  sudden- 
ly the  pressure  at  the  gear  teeth  is  so  high  and  so  suddenly 
applied  that  breakage  is  likely  to  occur.  To  prevent  this  a 
friction  clutch,  spring  coupling,  centrifugal  clutch  or  other 
equivalent  device  must  be  employed  between  the  engine  shaft 
and  the  rotor  wheels. 

One  solution  of  this  problem  is  shown  in  the  Sturtevant  super- 
charger, Fig.  309.  The  single-stage  blower  runs  at  10  times 
the  engine  speed  through  a  2  to  1  belt  drive  in  series  with  a  5  to 
1  helical  gear  drive  contained  within  the  blower  casing.  The  belt 
drive  is  vertical  and  is  brought  into  action  when  desired  by  an 

Induction  Pipes 


Bett  Drive 


Geared 
Blower. 


FIG.  309. — Sturtevant  supercharger. 

idler  pulley.  This  arrangement  gives  ample  slip  when  the 
engine  speed  charges  suddenly.  The  weight  of  the  super- 
charging device  in  this  case  is  stated  by  the  manufacturers  to  be 
50  Ib.  for  a  210-h.p.  engine.  The  engine  speed  with  constant- 
pitch  propeller  increases  from  about  2,100  r.p.m.  at  the  ground 
to  about  2,500  r.p.m.  at  20,000  ft.  altitude;  the  blower 
consequently  increases  from  21,000  to  25,000  r.p.m. 

An  English  design,  shown  in  Fig.  310,  has  a  double  reduction 
of  11  to  1  between  the  engine  shaft  and  the  blower  disc  with 
three  intermediate  wheels  distributing  the  torque  to  the  driven 
pinion;  the  over-all  efficiency  is  53  per  cent. 

German  constructions  show  multi-stage  compressors  with  the 
relatively  low  peripheral  speeds  of  400  to  500  ft.  per  second.1 
At  these  speeds  the  design  works  out  to  three  stages  to  maintain 

1  HILDESHEIM,  Automotive  Industries,  Oct.  21,  1920. 


SUPERCHARGING 


405 


406 


THE  AIRPLANE  ENGINE 


full  power  up  to  11,500  ft.  and  four  stages  to  16,000  ft.  The 
compressor  makes  10,000  to  11,000  r.p.m.  Such  compressors 
have  been  used  both  as  individual  superchargers  direct-coupled 
to  a  single  engine,  and  as  central  superchargers  driven  by  a 
separate  engine  and  delivering  compressed  air  to  all  the  engines 
of  a  multi-engine  plane.  With  a  separate  engine,  no  clutch  is 
necessary  between  engine  and  compressor  but  the  engine  must 
be  provided  with  a  flywheel  to  prevent  too  great  acceleration 
and  consequent  stripping  of  the  gear  wheels.  Individual  super- 
chargers are  generally  driven  from  the  rear  end  of  the  crankshaft, 


FIG.  311. — Schwade  multi-stage  centrifugal  supercharger. 

but  in  some  cases  the  gears  connect  to  the  propeller  end  of  the 
shaft  to  avoid  the  torsional  oscillations  which  have  sometimes 
given  much  trouble  at  the  rear  end. 

The  Schwade  three-stage  supercharger  shown  in  Fig.  311 
delivers  2,200  Ib.  of  air  per  hour  at  a  pressure  ratio  of  1  to  1.52 
(11,500  ft.  altitude).  The  shaft  speed  is  1,400,  intermediate 
gears  3,500,  blower  10,500  r.p.m.  The  rotor  diameter  is  10  in., 
peripheral  speed  460  ft.  per  second.  The  pinion  on  the  blower 
shaft  is  built  in  one  with  a  friction  clutch  consisting  of  four 
bronze  sectors  which  are  pressed  against  the  inside  of  the  clutch 
housing  by  centrifugal  force  and  which  come  into  action  when  the 
engine  speed  reaches  600  r.p.m.  The  casing  and  its  supports, 
partition  walls,  and  diffusors  are  of  aluminum.  The  super- 
charger complete  weighs  105  Ib.  and  was  applied  to  a  260-h.p. 


SUPERCHARGING 


407 


Mercedes  engine  weighing  925  lb.;  it  has  also  been  used  with 
rotary  engines.  A  four-stage  compressor  for  supercharging  up 
to  16,000  ft.  weighs  132  lb. 

Brown-Boveri  have  built  a  central  four-stage  supercharger, 
18.5  in.  diameter  and  making  6,000  r.p.m.,  which  gives  a  periph- 
eral speed  of  490  ft.  per  second.  The  gear  ratio  is  4.15  to  1. 
This  machine  supplies  9,200  lb.  of  air  per  hour  at  0.52  atmosphere 
initial  and  1  atmospheric  final  pressure;  it  is  driven  by  a  125-h.p. 
engine  and  sends  compressed  air  to  engines  aggregating  1,200 
h.p.  The  gear  teeth  are  of  the  Maag  form,  0.5  in.  circular  pitch, 


FIG.  312. — Spring  coupling  for  supercharger  drive. 

2  in.  face  width,  are  hardened  and  ground  and  are  loaded  to 
1,600  lb.  per  square  inch  at  full  load.  Oil  is  injected  directly 
between  the  teeth.  The  blower  is  connected  to  the  engine 
shaft  through  a  leather  block  joint.  The  coupling  has  a  disc 
flywheel  mounted  on  it  (weight  of  both  44  lb.)  to  give  smooth 
operation  and  protect  the  gears  against  shock.  Tests  with 
spring  couplings  have  shown  that  the  actual  forces  at  the  gear 
teeth  are  likely  to  be  four  times  the  normal  driving  force.  The 
details  of  a  successful  spring  coupling  are  shown  in  Fig.  312. 
The  efficiencies  of  these  multi-stage  compressors  (isothermal 
basis)  average  about  65  per  cent;  they  may  reach  68  per  cent  in 
some  cases. 


408 


THE  AIRPLANE  ENGINE 


With  a  supercharger  direct-coupled  to  the  engine  and 
operating  all  the  time,  a  throttle  valve  should  be  placed  on 
the  suction  side  of  the  supercharger — either  hand  or  auto- 
matically operated — to  keep  down  the  pressure  and  power  at  low 
altitudes.  With  this  location  of  the  throttle  valve  the  power 
absorbed  by  the  compressor  will  be  less  than  with  the  throttle 
on  the  discharge  side.  With  central  superchargers,  the  air 
pressure  is  controlled  by  throttling  the  supercharger  engine; 
the  air  pipe  to  each  engine  is  fitted  with  a  relief  valve,  a  throttle 

Carburetor 


Automatic  Air 
P?  Inlet  Va/ve 


Explosion  Valve 
js'5pray  Check 


FIG.  313. — Explosion  relief  valve. 

valve  for  the  compressed  air  and  an  automatic  air  admission 
valve  which  closes  when  the  supercharger  comes  into  action. 
Relief  or  explosion  valves  are  important;  without  such  a  valve 
a  back  fire  would  be  likely  to  destroy  the  partition  wall  between 
the  last  two  stages  of  the  blower.  Figure  313  shows  an 
explosion  valve  held  on  its  seat  by  springs  and  also  an  automatic 
inlet  valve  which  closes  whenever  the  compressor  is  in  use. 
The  difficulties  of  geared  drives  can  be  eliminated  and  greater 
total  power  obtained  by  the  use  of  an  exhaust  gas  turbine  for 
driving -the  compressor.  In  this  case  there  is  no  fixed  relation 
between  the  engine  and  compressor  speeds.  The  exhaust 


SUPERCHARGING 


409 


manifold  leads  to  the  nozzle  chamber  of  the  turbine  and  the 
compressor  is  mounted  on  the  turbine  shaft.  The  scheme  is 
shown  in  Fig.  314. 

Much  development  work  has  been  done  on  turbo-superchargers 


Intake  1o  Irrtute  ExhausHo  Turbine 

Engine  -.,  Va/ve 


Carburetor- 


:Turbine  Discharge 


"•  Air  Imps  Her 
Combined  Turbine 
and  Compressor  Sfrcrff- 

Air  Compressor 
Inlet 


Air  Discharge  to 
Induction  System 

FIG.  314. — Diagram  of  exhaust- turbine  supercharger. 

but  they  must  still  be  considered  as  in  the  experimental  stage. 
The  principal  difficulties  encountered  have  been  with  the  exhaust 
valves,  which  are  subjected  to  a  higher  temperature  and  which  are 
not  cooled  by  exposure  to  the  outside  temperatures;  with  the 
manifold,  which  is  kept  continuously  at  a  high  temperature  and 

r  n 


Turbine  Roto, 


Centrifugal 
Compresser 


FIG.  315. — Rateau  exhaust-turbine  supercharger. 

which  gives  expansion  troubles  and  difficulties  in  maintaining 
tight  joints;  and  with  the  nozzle  plate  and  blades  of  the  turbine, 
which  are  continuously  at  high  temperatures.  The  increase  in 
weight  due  to  the  supercharging  device  for  a  400-h.p.  engine  can 


410 


THE  AIRPLANE  ENGINE 


be  made  from  15  to  20  per  cent  of  the  engine  weight  when  a 
peripheral  speed  of  900  ft.  per  second  is  used  for  the  compressor. 
The  pioneer  work  on  turbo-superchargers  has  been  done  by 
Rateau  in  France.  Figure  315  shows  a  cross-section  of  his 
arrangement.  The  results  of  tests  of  the  Rateau  supercharger  at 


*     40 


|| 
I? 

t  e 

II 

£      70 


'20  30  40 

Pressure  in  Exhaust  Manifold, 
Inches  of  Mercury, Absolute. 

FIG.  316. — Pressures  in  Rateau  exhaust-turbine  supercharger. 

an  altitude  of  9,000  ft.  are  given  in  Figs.  316-318.  Figure  316 
shows  the  relation  between  the  back  pressure  on  the  engine  and 
the  pressure  at  the  carburetor,  the  exhaust  pressure  stays  at 
about  2  in.  of  mercury  above  the  carburetor  pressure.  Figure  317 
shows  the  variation  of  pressure  and  temperature  ratios  in  the 


cL 
I  1.1 


.0 


20000  22000  24000  26000  28000  30000 

R.p.m. 
FIG.  317. — Performance  of  Rateau  exhaust-turbine  supercharger. 

compressor  with  varying  r.p.m.  Figure  318  gives  the  variation 
of  over-all  efficiency  of  the  turbo  compressor  with  variation  of 
r.p.m.;  the  turbine  efficiency  naturally  increases  as  the  bucket 
speed  approaches  the  designed  speed.  Tests  of  a  Lorraine- 
Dietrich^S-cylinder,  160-h.p,  engine  show  an  increase  of  power 


SUPERCHARGING 


411 


from  111  to  164  h.p.  at  9,000  ft.  altitude  by  the  use  of  this  super- 
charger; the  engine  speed  increased  from  1,370  to  1,550  r.p.m. 
Tests  of  a  Breguet  plane  with  a  300-h.p.  Renault  engine  showed 
the  time  of  climb  to  16,400  ft.  decreased  from  47^  to  27  min.  and 
the  horizontal  speed  at  that  altitude  increased  from  91  to  120 
miles  per  hour  by  the  use  of  this  supercharger.  The  ceiling  was 
increased  13,000  ft.  and  the  speed  at  the  new  ceiling  was  25  per 
cent  greater  than  that  at  the  old  ceiling. 

British  tests  of  a  Rateau  supercharger  fitted  to  an  air-cooled 
engine  indicate  the  possibility  of  developing  within  12  per  cent 
of  ground  power  up  to  a  height  of  17,000  to  20,000  ft.  This  is 


0.300 

0.21& 
^  0.250 
I  0.225 
j|  0.200 
=  QMS 

0.150 

0.125 

0.100 

8000     10000  12000   14000   16000   18000  ?0000  22000 

Turbine  R.pm 
FIG.  318. — Over-all  efficiency  of  Rateau  exhaust-turbine  supercharger. 

obtained  by  maintaining  ground  pressure  at  the  carburetor, 
which  was  found  to  entail  a  back  pressure  of  about  19  Ib.  abs. 
at  the  exhaust. 

The  Moss  turbo-supercharger  (General  Electric  Co.)  has  been 
designed  for,  and  used  successfully  on,  the  Liberty  engine.  The 
exhaust  manifolds,  of  rectangular  form,  increase  in  cross-section 
as  they  come  forward  to  the  front  of  the  engine  and  join  at  the 
nozzle  box  (Fig.  319)  which  is  situated  inside  the  Vee  at  the  level 
of  the  tops  of  the  cylinders.  The  nozzles  cover  about  one-half 
the  circumference  of  the  wheel.  The  turbine  wheel  is  9.1  in.  in 
diameter,  the  compressor  10.5  in. ;  the  peripheral  speed  is  about 
1,000  ft.  per  second.  The  turbine  and  compressor  spindle  is 
supported  at  the  rear  in  a  water-cooled  bearing  mounted  on  the 
intake  pipe;  at  the  front,  the  bearing  is  in  the  air  intake  to  the 
compressor.  The  compressor  is  provided  with  guide  vanes 


/ 

/ 

/ 

/ 

/ 

/ 

/ 

/ 

/ 

/ 

/ 

/ 



^ 

^r 

412 


THE  AIRPLANE  ENGINE 


(Fig.  319)  and  discharges  at  the  bottom  into  the  intake  pipe 
which  extends  horizontally  backward  into  the  Vee  and  supports 
the  two  carburetors.  Performance  data  on  this  supercharger  are 
not  available  but  preliminary  tests  at  Pike's  Peak  (barometer 
18  in.)  showed  an  increase  of  engine  power  from  251  h.p.  to  367 
h.p.  when  running  at  1,800  r.p.m. 

The  use  of  a  supercharger  leads  to  some  complication  in  the 
fuel  supply  system.  The  carburetor  float  chamber  is  kept  at  the 
compressed-air  pressure,  which  is  variable,  and  may  be  7  or  8  Ib. 


Section  A- A  Section  B~ 


Exhausf  Manifold 


B 


X  I     *Exhau5l~Gas  Turbine 
\~Ceni-rifugctl  Blower 

-«— J 


FIG.  319. — Moss  exhaust- turbine  supercharger. 

per  square  inch  in  excess  of  the  atmospheric  pressure.  The  fuel 
has  to  be  fed  to  the  carburetor  against  this  pressure.  An  air 
pressure  system  is  undesirable  both  because  of  danger  of  leakage 
of  fuel  and  also  because  the  tanks  would  have  to  be  made  heavier 
to  withstand  such  high  pressures.  The  pressure  in  the  fuel  line 
must  not  exceed  the  pressure  in  the  float  chamber  by  an  amount 
sufficient  to  lift  the  float  valves  and  thereby  flood  the  carburetor; 
the  excess  of  pressure  should  be  less  than  5  Ib.  per  square  inch. 
A  practicable  system  is  to  supply  the  fuel  by  a  direct-acting 


SUPERCHARGING 


413 


engine-driven  fuel  pump  delivering  into  a  line  equipped  with  a 
spring-loaded  relief  valve  which  is  subjected  to  the  compressed- 
air  pressure  on  one  side  and  the  pump  discharge  pressure  on  the 
other.  The  spring  is  adjusted  to  lift  at  any  desired  excess  of 
fuel  over  air  pressure  and  by-passes  some  of  the  fuel  to  the  suction 
side  of  the  pump.  Such  a  valve  is  shown  in  Fig.  320.  The  air 
is  admitted  to  the  inside  of  a  corrugated  copper  "sylphon;" 
the  fuel  pressure  must  be 
sufficient  to  overcome  both 
the  air  pressure  and  the  spring 
compression. 

Another  method  is  to  main- 
tain the  supercharger  pressure 
in  the  gravity  tank  (Fig.  321) 
and  to  return  to  the  main 
tanks  the  excess  of  gasoline 
pumped  to  the  gravity  tank 
through  a  float-operated  valve. 

Small  fuel  tanks  with  air 
pressure  obtained  from  hand 
pumps  are  used  for  starting 
or  for  emergencies. 

Another  method  of  increas- 
ing the  power  of  an  airplane 
engine  at  high  altitudes  is  by 
supplying  the  engine  with  oxy- 
gen. Approximately  4  Ib. 

of  oxygen  is  necessary  to  burn  1  Ib.  of  gasoline.  This  weight 
is  so  considerable  that  oxygen  can  be  carried  only  for  emer- 
gency uses,  as,  for  example,  in  combat,  where  it  is  desired  to 
increase  the  speed  of  the  plane  for  a  short  time.  As  the  oxygen- 
gasoline  mixture  would  give  excessive  temperatures  and  pressures, 
the  oxygen  can  be  used  only  for  enriching  the  air  and  permitting 
a  moderate  increase  in  the  heat  developed  per  cycle.  The  oxygen 
must  be  carried  in  liquid  form  at  atmospheric  pressure,  otherwise 
the  weight  of  the  container  becomes  excessive.  As  the  boiling 
temperature  at  ground  pressure  of  liquid  oxygen  is  —  297°F.  it  is 
necessary  to  have  extraordinarily  good  heat  insulation  to  keep  the 
rate  of  evaporation  down  to  a  permissible  figure.  A  Dewar  flask 
is  the  only  practicable  insulator  in  this  case,  but  since  it  is  too 
fragile  and  brittle  to  withstand  the  slapping  of  the  liquid  in 


'From  Pump  D/scharge 


FIG.  320.— Relief  valve  for  fuel  line  of 
supercharged  engine. 


414 


THE  AIRPLANE  ENGINE 


flight,  it  is  necessary  to  put  the  liquid  oxygen  in  a  metal  container 
which  with  proper  cushioning  is  inserted  in  a  Dewar  flask.  The 
liquid  oxygen  can  be  evaporated  at  any  desired  rate  (1)  by 
electrical  heating  through  a  coil  immersed  in  the  liquid  and  (2)  by 
immersing  a  metal  rod  in  the  liquid  to  an  adjustable  depth  and 
utilizing  the  thermal  conduction  along  the  rod,  or  (3)  the  liquid 
can  be  siphoned  over  into  the  charge  going  to  the  cylinder.  It  is 
evident  that  special  adjustment  of  the  carburetor  is  necessary  to 
meet  the  condition  of  oxygen  supply. 


Gravity 
Tank- 


Pressure  Compensatt'nq  Pipe 


Float  for 

Operating 

Valves 


Carburetor 


From  the 
Supercharger 


TTTTTF 


To  The  Fue!  Tanks 


FIG.  321. — Fuel  system  for  supercharged  engine. 


The  power  developed  in  a  supercharged  engine  can  be  absorbed 
satisfactorily  only  by  the  use  of  a  variable  pitch  propeller.  A 
four-bladed  propeller  would  be  best  in  order  to  avoid  excessive 
peripheral  speeds  at  high  altitudes.  The  supercharger  may  be 
used  either  (a)  to  maintain  constant  power,  in  which  case  the 
revolutions  per  minute  will  vary  inversly  as  the  cube  root  of  the 
air  density,  and  the  torque  will  vary  as  the  cube  of  the  density 


SUPERCHARGING 


415 


or  (6)  to  maintain  constant  torque,  in  which  case  the  revolutions 
per  minute  and  power  will  vary  inversely  as  the  density. 

A  calculation  for  a  machine  of  standard  type  with  a  ground 
speed  of  110  miles  per  hour  gives  the  following  results: 


^ 

Altitude, 

10,000  ft. 

Altitude, 

20,000  ft. 

Type  of  engine 

Maximum 
speed, 
miles  per 
hour 

Rate  of 
climb, 
feet  per 
minute 

Maximum 
speed, 
miles  per 
hour 

Rate  of 
climb, 
feet  per 
minute 

Ceiling, 
feet 

105 

450 

19  000 

Constant  power  
Constant  torque 

, 
126 
132 

1,050 
1,420 

141 
159 

930 
1  770 

61,000 
No  theo- 

retical limit 

CHAPTER  XVII 


MANIFOLDS  AND  MUFFLERS 

Air  Intakes. — The  location,  dimensions  and  orientation  of 
the  air  intake  to  the  carburetor  have  considerable  influence  on 
the  capacity  of  the  engine.  Attempts  are  often  made  to  increase 
the  density  of  the  air  going  to  the  carburetor  by  having  the  air 
intake  face  full  forward,  so  as  to  establish  in  the  intake  a  pressure 
which  is  the  sum  of  the  surrounding  atmospheric  pressure  and  the 
velocity  head  equivalent  to  the  air  velocity  of  the  plane  or  of 
the  slip  stream.  With  a  plane  making  120  miles  per  hour  at  the 
ground  this  dynamic  head  would  be  about  0.24  Ib.  per  square 
inch.  The  necessity  for  keeping  the  air  intake  pipe  of  moderate 
length  does  not  usually  permit  the  intake  to  be  located  in  free  air  so 
as  to  take  advantage  of  this  velocity  head.  Inside  the  fuselage 
there  is  very  small  air  velocity  and  the  practical  means  of  getting 
access  to  high  velocity  air  is  by  extending  the  air  intake  pipe 

above  the  cowling.  This  arrange- 
ment has  the  great  advantage 
that  spillage  of  gasoline  into  the 
fuselage  is  thereby  prevented  and 
a  frequent  source  of  fires  is  elimi- 
nated. The  disadvantage  of  this 
arrangement  when  applied  to  the 
usual  form  of  carburetor  is  that  the 
air  intake  pipe  will  have  two  or 
more  right-angle  turns.  With  an 
inverted  type  of  carburetor,  with 
jets  discharging  downward  this  dif- 
nculty  is  overcome  and  a  short  di- 
rect air  intake  pipe  can  be  employed. 
Investigations  in  England  on  the  best  shape  of  air  intake 
have  shown  that  maximum  power  is  obtained  when  the  intake 
pipe  is  cut  off  at  an  angle  of  25  deg.  to  the  horizontal  so  as  to 
form  a  scoop  and  that  instead  of  mounting  the  scoops  dead  ahead 
they  are  best  turned  about  7  deg.  to  compensate  for  the  propeller 
slip  stream  (see  Fig.  322).  The  double  right-angle  turn  of  the 

416 


'••Cowling 
•;Carburefor 


FIG.    322. — British   design  for 
intake  pipe. 


MANIFOLDS  AND  MUFFLERS 


417 


•25° 


Stand 
Pipe 


air  on  its  way  to  the  carburetor  results  in  a  disturbed  air  flow 
which  can  be  largely  corrected  by  the  insertion  of  the  baffle  plate 
indicated.  In  the  Liberty  engine  a  single  standpipe  has  been 
used  to  serve  two  carburetors,  by  the  use  of  the  duplex  air  intake 
shown  in  Fig.  323. 

The  length  of  intake  pipe  will  affect  the  engine  capacity 
because  the  inertia  of  the  air  column  tends  to  make  the  flow  to  the 
carburetor  continuous  in  spite  of 
the  intermittent  action  of  the 
cylinder  suctions.  In  the  Lib- 
erty engine  the  optimum  length 
with  inverted  type  of  carburetor 
is  found  to  be  6  in.  at  normal 
revolutions  per  minute. 

Intake  Manifolds. — The  main 
function  of  the  intake  manifold 
is  the  distribution  of  the  mixture 
formed  in  a  carburetor  to  several  FlG-  323.— Liberty  engine  duplex  in- 

v     j  -n  i        £CL    •  take  Plpe' 

cylinders.     For  good  efficiency 

and  capacity  the  strength  and  density  of  the  mixture  reaching 
all  cylinders  should  be  the  same.  To  ensure  these  results  the 
amount  of  fuel  entering,  and  the  pressure  drop  from  carburetor 
to  inlet  valve,  should  be  the  same  in  all  branches.  With  a  fuel 
which  is  not  completely  vaporized  before  the  branching  of  the 
manifold  occurs  it  is  very  difficult,  if  not  impossible,  to  ensure 
proper  distribution  of  the  fuel.  This  difficulty  increases  as  the 
volatility  of  the  fuel  decreases  and  is  therefore  greater  with 
ordinary  commercial  gasoline  than  with  airplane  fuels. 

For  the  vaporization  of  a  volatile  fuel  the  important  factors 
are  (1)  air  supply  of  sufficiently  high  temperature,  (2)  fine 
atomization  at  the  jet,  (3)  avoidance  of  obstacles  in  the  path  of 
the  mixture  so  that  the  atomized  liquid  drops  may  have  no 
opportunity  for  coalescence  before  being  vaporized,  and  (4) 
sufficient  time  for  the  vaporization.  The  temperature  of  the 
mixture  after  vaporization  is  much  below  that  of  the  entering 
air  because  the  latent  heat  of  vaporization  of  the  fuel  is  taken 
from  the  air.  Assuming  a  latent  heat  of  135  B.t.u.  per  pound, 
complete  vaporization  will  produce  a  temperature  drop  of 
47°F.  with  an  air  fuel  ratio  of  10  to  1,  and  a  drop  of  25°F.  with 
an  air  fuel  ratio  of  20  to  1.  If  all  the  fuel  is  not  vaporized  the 
temperature  drop  will  be  less.  The  fall  of  temperature  in  the 

27 


418 


THE  AIRPLANE  ENGINE 


manifold  may  bring  about  a  deposit  of  ice  both  inside  and 
outside  the  manifold  as  a  result  of  the  freezing  of  the  moisture 
in  the  air.  It  has  been  suggested1  that  accumulation  of  ice 
inside  the  manifold  may  be  the  cause  of  numerous  unexplained 
engine  failures  and  resulting  crashes.  The  relation  between  the 
manifold  temperature,  measured  at  the  intake  valve,  and  the 
air  supply  temperature  in  a  Liberty  engine  is  shown  in  Fig.  324. 
The  manifold  was  water-jacketed.  Below  20°F.  there  was  so 
little  evaporation  that  the  manifold  temperature  was  higher 
than  the  air  supply  temperature.  As  the  air  supply  temperature 
increased  up  to  120°F.  the  temperature  fall  increased  up  to 
about  35°F.  The  air  fuel  ratio  was  about  16  to  1. 


i-'  rco 

_c 

& 

CT>  100 

£ 
€  so 

-t- 
c 

SL  6° 
B 

£    40 
33 
o 

n 

D 

:E 


0        20      40       00      80       100     120 
Air  Intake  Tern perature,Dec}  Fahr. 
FIG.  324. — Temperature  change  in  intake  manifold. 

With  air  initially  cold  it  will  not  be  possible  to  vaporize  the 
fuel  completely  by  heat  absorbed  from  the  air  because  at  low 
temperatures  the  air  will  become  saturated  with  the  fuel  vapor 
before  all  the  fuel  is  vaporized.  Table  14,  page  229,  shows,  for 
example,  that  with  a  theoretically  correct  mixture,  the  air  cannot 
hold  all  the  pentane  in  the  vapor  form  at  a  temperature  below 
about  38°F.  In  such  case  the  only  chance  for  complete  vapor- 
ization is  by  supplying  heat  from  outside.  This  is  accomplished 
by  utilizing  some  of  the  heat  either  of  the  jacket  water  or  of  the 
exhaust  gases.  The  possibilities  are  (1)  to  preheat  the  air  before 
it  enters  the  carburetor,  (2)  to  heat  the  mixture  in  the  manifold, 
and  (3)  to  heat  the  manifold  locally  at  some  place  on  which  the 
1  SPARROW,  Technical  Note,  No.  55,  Nat.  Adv.  Comm.  Aeronautics. 


MANIFOLDS  AND  MUFFLERS 


419 


liquid  drops  impinge  so  as  to  supply  heat  to  the  liquid  only 
(hot-spot  method). 

All  preheating  is  objectionable  in  that  it  diminishes  the  density 
of  the  charge  and  thereby  decreases  the  capacity  of  the  engine. 
That  method  of  preheating  is  best  which  causes  vaporization 
with  the  minimum  resulting  temperature  of  the  mixture.  All 
three  methods  of  preheating  are  employed  in  airplane  practice. 
In  some  of  the  German  engines  preheating  the  air  is  accomplished 
by  taking  the  air  through  pipes  in  the  crankcase  (Fig.  78),  which 
has  the  advantage  of  cooling  the  lubricating  oil.  The  more 
common  procedure  is  heating  the  manifold  by  jacket  water. 


FIG.  325. — Liberty  engine  intake  manifold. 

There  is  no  general  consensus  of  opinion  as  to  the  best  form  of 
manifold.  It  is  desirable  that  sharp  turns  should  be  avoided 
as  far  as  possible,  that  the  various  branches  should  have  approxi- 
mately the  same  length,  that  sudden  enlargements  should  be 
avoided  and  that  the  velocities  should  be  high  enough  to  prevent 
deposition  of  liquid  drops  but  not  so  high  as  to  cause  a  large 
frictional  resistance.  Mean  velocities  of  120  to  200  ft.  per  minute 
are  common  but  values  up  to  250  ft.  per  minute  have  been  used 
successfully. 

Manifolds  usually  divide  themselves  into  two  classes,  the 
short-branch  type  and  the  long-branch  type.  The  standard 
Liberty  manifold  (Fig.  325)  is  a  good  example  of  the  short-branch 


420 


THE  AIRPLANE  ENGINE 


type;  it  is  water-jacketed  and  has  a  baffle  plate  opposite  to  the 
inlet  to  equalize  the  lengths  of  the  three  branches.  The  Benz 
manifold  (Fig.  326)  is  an  example  of  an  unjacketed  long-branch 
manifold  with  all  three  branches  of  the  same  length  and  with  long- 
turn  elbows.  Another  design  for  accomplishing  the  same  purpose 
is  shown  in  the  Hall-Scott  engine  (Fig.  63). 

Manifolds  for  vertical  engines  can  usually  be  arranged  in 
any  way  the  designer  likes.  In  the  Maybach  engine  (Fig.  80) 
the  carburetors  are  at  the  ends  and  the  manifolds  run  along  the 
side  of  the  engine  with  no  attempt  to  equalize  the  lengths  of  the 
branches.  Ordinarily  such  arrangements  as  those  of  Figs.  325  and 
326  are  used. 

In  90-deg.  Vee  engines  there  is  plenty  of  room  in  the  Vee  to 
accommodate  the  carburetors  and  the  intake  is  usually  inside  the 


FIG.  326. — Benz  intake  manifold. 

Vee.  With  this  location  a  short-branch  manifold  must  be  used. 
The  Hispano-Suiza  engine  (Fig.  50)  shows  a  typical  arrange- 
ment with  the  transverse  pipe  water-jacketed.  If  long-branch 
manifolds  are  to  be  used  they  must  either  be  placed  in  the  rear  of 
the  engine,  as  in  the  Curtiss  engines,  Figs.  59  and  62,  or  the 
intake  valves  must  be  on  the  outside  of  the  Vee. 

In  60-deg.  and  45-deg.  Vee  engines  the  space  inside  the  Vee  is 
small  and  a  favorable  design  of  manifold  is  difficult  if  the  car- 
buretors are  placed  inside  the  Vee.  If  the  inlet  valves  are  placed 
outside  the  Vee,  the  exhaust  pipes  will  be  crowded  inside  the  Vee 
and  may  give  rise  to  troubles  caused  by  their  proximity  to  the 
valve  springs,  etc.  An  alternative  arrangement  is  to  provide  a 
space  between  the  two  central  cylinders  of  each  block  (as  in  the 
Bugatti  (Fig.  67)  and  Fiat  engines  (Fig.  76))  and  to  lead  the 
induction  pipes  from  carburetors  mounted  outside  the  Vee 
through  these  spaces  to  manifolds  inside  the  Vee. 


MANIFOLDS  AND  MUFFLERS  421 

In  radial  and  rotary  engines  the  distribution  problem  is  com- 
paratively simple,  especially  where  an  induction  chamber  is 
provided  in  the  crankcase.  The  special  distributing  chamber  of 
the  Bristol  " Jupiter"  engine  (Fig.  148)  is  noteworthy. 

Exhaust  Manifolds. — The  function  of  the  exhaust  manifold  is, 
primarily,  to  conduct  the  exhaust  gases  away  from  the  airplane 
with  the  minimum  back  pressure  at  the  engine  and  without  fire 
risk  to  the  airplane  or  annoyance  from  the  discharged  gases  to 
the  pilot.  An  additional  function  may  be  to  muffle  the  sound  of 
the  exhaust,  although  this  has  usually  been  considered  unimpor- 
tant in  military  machines.  The  manifold  is  usually  required  to 
have  a  clearance  of  2^  in.  from  wooden  parts  and  of  3H  m-  from 
fabric  parts  of  the  airplane. 


FIG.  327. — Hall-Scott  exhaust  manifold. 

The  simplest  exhaust  piping  is  either  its  complete  absence 
as  in  rotary  engines  and  certain  stationary  engines  or  the  use  of 
short  tubes  discharging  outwards  or  upwards.  In  some  engines 
these  stub  tubes  are  cut  off  on  the  outer  end  at  an  angle  of  45 
deg.  so  as  to  discharge  backwards  as  well  as  outwards.  The 
absence  of  exhaust  pipes  or  the  use  of  short  straight  pipes  is 
advantageous  not  only  in  reducing  back  pressure  but  also  in 
permitting  better  cooling  of  the  exhaust  valves  by  radiation  and 
avoiding  heating  of  the  exhaust  valve  springs.  This  arrangement 
does  not  conduct  the  gases  away  from  the  crew  of  the  airplane. 

A  method  of  discharging  the  gases  overhead  and  to  the  rear 
with  small  back  pressure  is  shown  in  the  Hall-Scott  manifold  of 
Fig.  327.  This  arrangement  obstructs  the  view  of  the  pilot  in  a 
tractor  machine  but  diminishes  the  noise  heard  from  below. 
Long  radius  curves  are  very  essential  for  all  the  branches  if  back 


422  THE  AIRPLANE  ENGINE 

pressure  is  to  be  kept  down;  the  radius  should  be  about  2K  times 
the  diameter  of  the  pipe.  The  arrangement  of  Fig.  328  with 
discharge  to  the  rear  and  with  an  exhaust  main  of  increasing 
diameter  not  only  carries  the  gases  away  from  the  crew  but  slows 
down  the  gases  before  exit  and  thereby  tends  to  diminish  noise. 
As  the  exhaust  period  lasts  more  than  two-thirds  of  a  revolution 
there  is  overlapping  of  exhausts  in  a  manifold  connecting  three  or 
more  cylinders.  The  velocity  of  the  gases  immediately  after  the 
opening  of  the  exhaust  valve  is  extremely  high  and  its  effect  on 
the  exhaust  from  any  other  cylinder  whose  exhaust  valve  is  open 
at  that  time  should  be  carefully  considered.  By  the  use  of  two 
concentric  exhaust  mains  to  which  the  cylinders  are  connected  so 


FIG.  328. — Hall-Scott  exhaust  manifold. 

that  no  two  consecutive  exhausts  go  into  the  same  main,  an 
ejector  effect  can  be  obtained  from  the  action  of  a  newly  opened 
exhaust  on  the  exhaust  which  is  closing,  which  may  cause  substan- 
tial scavenging  in  the  closing  cylinder. 

The  cross-section  area  of  manifold  branches  is  governed  by  the 
size  of  the  ports  in  the  cylinders;  an  average  value  is  about  0.15 
sq.  in.  per  brake  horse  power  of  the  cylinder. 

Mufflers,  to  be  efficient,  must  slow  down  the  exhaust  gases  to 
velocities  below  that  of  sound  (1,100  ft.  per  second);  actual  gas 
velocities  probably  exceed  2,000  ft.  per  second.  For  airplane  use 
the  muffler  must  be  of  light  weight  and  yet  durable.  The  con- 
structions employing  reversal  of  direction  of  the  gases  and  dis- 
charge through  small  holes  are  likely  to  give  excessive  back 
pressure.  Tests  by  Diederichs  and  Upton  show  that  the  volume 
of  the  muffler  should  be  about  three  times  that  of  a  single  cylinder 


MANIFOLDS  AND  MUFFLERS 


423 


of  the  engine,  and  that  the  inlet  to  the  muffler  should  be  tangen- 
tial so  as  to  give  the  gas  a  whirling  motion.  For  durability  the 
muffler  should  be  attached  to  the  end  of  a  tail  pipe  6  or  8  ft.  long 
which  will  cool  the  gases  sufficiently  to  prevent  excessive  oxida- 
tion of  the  muffler.  One  of  the  simplest  and  most  successful 
mufflers  is  shown  in  Fig.  329.  The  tangential  inlet  pipe  starts  of 
circular  cross-section  and  flattens  out  in  a  fan  shape  so  as  to  give 
admission  for  almost  the  whole  length  of  the  muffler.  An  inner 


Inlef 


Inlet 


FIG.  329. — Diederichs  and  Upton  exhaust  muffler. 

shell,  AB,  is  provided  with  a  large  number  of  holes  except  in  the 
small  arc  between  A  and  B;  these  holes  are  smallest  near  B  and 
increase  in  diameter  from  B  to  A  (contra-clockwise).  The  gas 
passes  through  these  holes  and  then  through  more  holes  in  the 
innermost  shell  and  finally  escapes  at  the  open  end  of  the  inner- 
most shell.  With  this  muffler  the  exhaust  noise  can  be  reduced 
80  per  cent  with  about  two-thirds  of  1  per  cent  reduction  in 
engine  power.  Mufflers  of  this  type  are  found  to  give  less  back 
pressure  than  those  with  axial  admission. 


CHAPTER  XVIII 
STARTING 

The  starting  of  an  airplane  engine  depends  on  three  things, 
(1)  obtaining  an  explosive  mixture  in  the  cylinder,  (2)  a  device 
for  igniting  it,  and  (3)  a  device  for  turning  the  engine  over. 

Hand  Starting. — The  idling  device  on  the  carburetor  will  fur- 
nish a  rich  mixture  to  the  cylinders  if  the  throttle  valve  is  closed 
and  the  engine  is  turned  at  a  sufficient  speed.  If  the  revolutions 
per  minute  of  the  engine  is  too  low  (below  20  to  30  r.p.m.)  the 
velocity  of  the  air  in  the  intake  manifold  will  be  insufficient  to 
carry  the  fuel  into  the  cylinders;  this  will  be  the  usual  condition 


FIG.  330. — Priming  system  for  a  12-cylinder  Vee  engine. 

with  hand  starting.  To  overcome  this  difficulty  the  cylinders  may 
be  primed  through  individual  priming  cocks  or  through  such  a 
priming  system  as  is  shown  in  Fig.  330.  With  the  engine  cold 
only  a  small  part  of  the  gasoline  will  be  vaporized  and  the  rest 
will  go  as  liquid  into  the  cylinder  and  will  dilute  the  lubricant. 
In  very  cold  weather  there  will  be  difficulty  in  vaporizing  enough 
of  the  gasoline  to  form  an  explosive  mixture  unless  the  jacket  is 
filled  with  hot  water  or  some  other  device  has  been  used  to  heat 
the  engine.  In  such  a  case  a  number  of  attempts  may  have  to  be 

424 


STARTING  425 

made  before  the  engine  starts,  and  as  an  excess  of  gasoline  is  put 
in  the  cylinder  before  each  attempt  the  walls  may  be  washed 
clear  of  lubricant  and  scoring  of  the  cylinder  may  occur  when 
the  engine  finally  starts. 

To  overcome  these  difficulties  a  more  volatile  fuel  such  as 
ether  may  be  used  for  priming.  A  more  satisfactory  practice, 
for  cold  weather,  is  to  use  hydrogen  as  the  starting  fuel.  This 
has  the  great  advantage  that  it  does  not  have  to  be  vaporized  and 
that  it  forms  an  explosive  mixture  throughout  a  great  range  of 
strengths.  A  hydrogen-air  mixture  will  explode  if  the  hydrogen 
forms  from  10  to  66  per  cent  of  the  mixture  by  volume;  with 
gasoline  vapor  the  range  is  only  from  1.5  to  4.8  per  cent.  The 
hydrogen  may  be  admitted  through  the  priming  system  as 
shown  in  Fig.  330 ;  it  is  best  to  take  the  gas  from  a  fabric  balloon 
in  which  it  is  at  atmospheric  pressure  and  not  from  a  high-pres- 
sure bottle  which  would  be  likely  to  give  an  excess  of  gas.  The 
hydrogen  must  be  shut  off  shortly  after  the  starting  as  it  gives 
more  violent  explosions  than  gasoline. 

The  regular  magneto  will  not  turn  fast  enough  with  hand 
starting  to  give  a  spark.  If  a  battery  system  of  ignition  is  used 
there  is  no  difficulty  from  this  source.  If  a  magneto  system  is 
used  it  is  customary  to  supply  a  hand-operated  starting  magneto 
which  is  geared  to  run  at  high  speed  and  is  turned  independently 
of  the  engine.  With  hand  starting  the  engine  may  be  pulled 
over  several  times  to  fill  the  cylinders  with  explosive  mixture, 
after  which  a  shower  of  sparks  is  sent  from  the  starting  magneto 
through  the  distributor  to  the  fully  retarded  cylinder  which  is 
ready  to  fire,  or  the  ignition  may  be  left  on,  fully  retarded,  while 
the  propeller  is  pulled  over  either  by  hand  or  by  rope. 

The  directions  for  starting  the  Liberty  engine  by  hand  are  as 
follows:  Inject  jj^>  oz.  of  lubricating  oil  through  each  priming 
cock.  Turn  switch  "off."  Turn  the  engine  over  five  times. 
Open  throttle  slightly.  Retard  spark  fully.  Prime  each  cylinder 
twice.  Turn  engine  over  twice.  Turn  on  one  switch.  Pull 
down  and  forward  on  propeller  blade.  After  the  engine  starts: 
Advance  spark  half  way.  Turn  on  both  switches.  Leave 
throttle  undisturbed  for  5  min.  The  lubricating  oil  should  be 
warm  by  the  time  the  jacket  outlet  water  has  reached  150°. 
If,  in  cold  weather,  it  is  not,  stop  the  engine  for  5  min. ;  then  start 
all  over  again.  Accelerate  and  slow  down  the  engine  a  few 
times.  Note  that  the  oil  gage  registers  pressure,  5  Ib.  at  600 


426 


THE  AIRPLANE  ENGINE 


r.p.m.  At  this  speed,  the  ammeter  should  show  "  discharge.7' 
At  1,000  r.p.m.  it  should  indicate  "  charge."  After  5  min.  more, 
open  the  throttle  wide.  The  speed  should  rise  to  about  1,600 
r.p.m. 

The  necessity  of  turning  the  engine  over  preliminary  to 
starting  can  be  avoided  by  the  use  of  a  mechanism  which  will 
lift  the  valves  and  permit  the  pumping  of  an  explosive  charge 
into  the  cylinders.  In  the  Maybach  engine  this  is  accomplished 

(Fig.  331)  by  lifting  all  the 
tappets  (inlet  and  exhaust) 
off  their  cams  by  depressing 
the  hand  lever,  A,  which 
rotates  the  two  lay  shafts, 
BB,  and  lifts  the  tappets 
through  slots  in  the  lay 
shafts  engaging  with  small 
lugs  projecting  from  the  tap- 
pets. At  the  same  time  the 
shutter,  C,  in  the  exhaust 
main  is  closed.  The  hand 
pump,  E,  is  then  operated 
and  air  is  sucked  through  the 
carburetor  and  the  intake 
manifold  and  through  the 
cylinder  to  the  exhaust  man- 
ifold and  to  the  pump.  When 
the  cylinders  are  thus  charged 
the  lever,  A,  is  brought  ver- 
tical, which  restores  the  engine  to  its  operating  position.  The 
starting  magneto  is  then  used  to  start  the  engine.  The  whole 
operation  can  be  carried  out  from  the  pilot's  seat. 

The  starting  torque  in  high-power  airplane  engines  is  yery 
considerable.  The  load  consists  of  two  parts,  (1)  the  compression 
load,  and  (2)  the  friction  load.  The  compression  load  increases 
with  the  cylinder  diameter  and  the  compression  ratio.  The 
maximum  torque  does  not  change  with  the  number  of  cylinders 
because  high-compression  pressure  will  exist  at  any  moment  in 
one  cylinder  only.  As  the  compressed  gas  re-expands  the  mean 
torque  for  two  revolutions  does  not  increase  with  increase  in 
number  of  cylinders.  Friction  is  the  principal  resistance  in 
starting  the  engine,  especially  in  cold  weather  when  the  lubricating 


FIG.  331. — Starting  mechanism  of  May- 
bach  engine. 


STARTING 


427 


oil  may  be  near  the  solidifying  point.     Tests  at  McCook  Field1 

to  determine  the  average  starting  torque  for  various  engines  have 

yielded  the  following  results.     The  average  air  temperature  was 

AVERAGE  STARTING  TORQUE,  POUNDS-FEET 


Engine 

Number  cylinders 

Rated  horse  power 

Normal  speed 

t 

| 

g" 
1 

Stroke,  inches 

Number  of  engines 
tested 

Number  of  trials 

Starting  torque, 
pounds-feet 

Throttle 
open 

Throttle 
closed 

Maxi- 
mum 

l» 

H 

J» 

Liberty 

12 
8 
6 
8 
6 

400 
300 
210 
180 
185 

,700 
,800 
,700 
,600 
,400 

5.00 
5.51 
5.00 
4.75 
5.90 

7.00 
5.91 
7.00 
5.25 
7.08 

2 
2 
2 
1 
1 

.6 
6 
6 
3 
3 

130 
106 
133 
110 
150 

124 
102 
105 
82 
139 

156 
101 
135 
87 
153 

143 
96 
110 
77 
145 

Hispano-Suiza  
Liberty 

Packard 

BMW 

75°F.  With  freezing  temperatures  the  results  would  have  been 
much  higher — probably  doubled.  It  should  be  noted  that  the 
starting  torque  does  not  increase  nearly  as  rapidly  as  the  engine 
horse  power. 


FIG.  332. — Compression  release  of  Benz  engine. 

From  the  preceding  table  it  may  be  presumed  that  with  low 
temperature  the  mean  starting  torque  for  a  400-h.p.  engine  will 
be  from  300  to  400  Ib.-ft.  It  is  difficult  for  a  mechanic  to  exert 
this  torque  on  the  propeller  even  in  a  land  plane  without  danger  to 
himself  from  overbalancing;  in  a  sea  plane  it  is  even  more  difficult 
to  accomplish.  The  starting  torque  can  be  diminished  by  reduc- 
ing or  eliminating  the  compression.  For  this  purpose  compres- 
sion release  cams  may  be  provided  on  the  camshaft  to  keep  the 
exhaust  valves  open  during  the  compression  stroke.  In  the 
Benz  engine  (Fig.  332)  the  compression  release  cams  are  brought 

1  Air  Service  Information  Circular,  No.  126. 


428 


THE  AIRPLANE  ENGINE 


FIG.  333. — Compression  release  of  Basse- 
Selve  engine. 


into  action  by  the  axial  movement  of  the  camshaft  effected  by  a 
square-thread  screw  operated  by  a  small  lever  at  the  rear  of  the 
crankcase.  The  camshaft  is  returned  to  its  normal  running 
position  by  a  spring  inside  the  front  end  of  the  shaft.  The  relief 
cams  open  the  exhaust  valves  35  deg.  early  and  close  them  at 
22  deg.  late. 

In  the  Basse-Selve  engine  (Fig.  333)  the  compression  release  is 

operated  by  means  of  a  rod 
which  lies  horizontally  along 
the  outside  of  the  camshaft 
casing  directly  underneath 
the  exhaust-valve  rocker 
arms.  The  rod  is  slotted  in 
such  a  way  (see  separate 
detail,  Fig.  333)  as  to  form 
cams  which  lift  the  exhaust 
valve  rockers  when  the  rod 
is  partially  rotated  by  means 
of  the  hand  lever  at  the  end  of  the  rod.  With  this  device  the 
exhaust  valves  remain  open  so  long  as  the  rod  is  in  the  rotated 
position. 

In  addition  to  the  difficulty  which  is  experienced  in  starting 
large  engines  by  the  propeller  there  is  a  considerable  element  of 
danger,  which  has  proved  fatal  in  many  cases.  To  obviate  this, 
recourse  may  be  had  to  a  portable  engine  cranker,  which  is  usually 
available  only  at  airdromes,  or  to  a  starting  mechanism  integral 
with  the  engine.  Whatever  the  nature  of  the  starting  mechanism 
it  should  be  thrown  automatically  out  of  action  as  soon  as  the 
engine  fires  and  it  should  also  go  out  of  action  if  the  engine  fires 
before  the  dead  center  and  starts  to  turn  backward.  The  con- 
nection from  the  starter  to  the  engine  is 
through  a  dog  or  clutch,  as  in  Fig.  334, 
which  is  pushed  out  of  mesh  when  the 
engine  starts  forward  but  will  remain  in 
mesh  if  the  engine  starts  backward.  FlG-  334>~ Dog  clutch' 

The  acceleration  of  the  engine  shaft  as  a  result  of  firing  one 
of  the  cylinders  is  very  great  and  if  the  engine  starts  backward 
this  acceleration  will  be  transmitted  to  the  starting  mechan- 
ism. As  the  starting  mechanism  is  always  geared,  with  a  ratio 
which  may  be  100:1  or  more,  the  motivating  element  will  be  given 
an  enormous  acceleration  which  will  produce  high  teeth  pressures 


STARTING 


429 


in  the  gears  and  will  probably  strip  the  teeth  of  the  gears  or  the 
dog  projections  of  the  clutch.  If  the  gears  are  hand-operated 
through  a  crank  the  gear  ratio  may  be  10  or  20  to  1,  and,  although 
the  teeth  may  hold  on  a  back  fire,  the  mechanic  will  be  endan- 
gered by  the  high  speed  of  the  handle.  To  safeguard  the  starting 
gear  some  kind  of  friction  clutch  or  other  safety  drive  should  be 
placed  close  to  the  dog.  Multiple-disc  clutches  are  suitable  on 
account  of  their  compactness  and  low  weight,  but  all  such  friction 
devices  are  uncertain  in  their  action. 

A  different  type  of  safety  device  with  an  oscillatinggmember 
is  shown  in  Fig.  335.  \  The  oscillating 
member,  A,  takes  no  part  in  the  trans- 
mission of  the  load.  The  engine  drive 
is  to  the  right  of  the  figure.  In  normal 
rotation,  A  is  driven  by  the  gear  wheel, 
C,  and  is  free  to  oscillate  as  determined 
by  the  tapered  sides  of  the  stationary 
projections  on  D.  If  a  reverse  rotation  FlG-  335.— Safety  clutch, 
occurs  D  prevents  the  rotation  of  A,  which  causes  the  sleeve 
on  which  C  is  mounted  to  move  to  the  right,  and  thereby  throws 
C  out  of  mesh.  As  the  action  cannot  take  place  until  some  back- 
ward movement  has  occurred  this  gear  should  be  placed  on  a 
shaft  geared  to  the  engine  shaft,  so  that  release  may  occur 
quickly.  In  a  hand-operated  system  it  might  be  on  the  handle 
shaft. 

Integral  cranking  mechanisms  are  either  operated  by  hand, 
by  compressed  air,  or  by  electric  motor.  It  is  usually  necessary 
to  have  the  engine  rotating  at  a  speed  of  from  10  to  20  r.p.m. 
before  regular  operation  will  take  place.  Hand  mechanism  must 
have  the  operating  handle  at  the  side  or  rear  of  the  engine  and 
consequently  must  employ  worm,  bevel,  or  helical  gears,  usually 
in  conjunction  with  spur  gears.  The  efficiency  of  the  gear  train 
is  poor  and  the  attainable  cranking  speed  very  low.  The  hand 
mechanism  for  the  Hispano-Suiza  engine  is  shown  in  Fig.  336. 
It  includes  double-reduction  spur  gears,  a  dog  clutch,  a  releasing 
spring  and  a  gear-driven  starting  magneto. 

Compressed  air  may  be  utilized  (1)  in  a  motor  which  cranks 
the  engine,  (2)  it  may  be  carbureted  and  sent  through  a  dis- 
tributor and  exploded  in  the  cylinders,  or  (3)  it  may  go  direct 
to  the  cylinders  at  a  high  pressure  through  a  distributor  and 

SHERMAN,  The  Automobile  Engineer,  December,  1919. 


430 


THE  AIRPLANE  ENGINE 


special  non-return  valves.     If  a  compressed  air  motor  is  used,  it 
can  be  run  from  the  engine  as  an  air  compressor  to  store  up  the 


-'Starting 
Magneto 


FIG.  336. — Hand-starting  mechanism  for  Hispano-Suiza  engine. 


-  me  rear 
flange,   the 


-Engine  Face 


FIG.  337. — Radial  air  compressor. 

air  necessary  for  starting.     In  the  Motor-Compressor  Company's 
starter  (Fig.  337)  a  multi-cylinder  radial  compressor  is  lined  up 


STARTING 


431 


with  the  crankshaft  at  the  rear  of  the  engine.  When  operated 
as  a  compressor  it  is  directly  connected  through  a  positive  clutch 
to  the  engine  and  runs  at  engine  speed,  compressing  air  up  to 
230  Ib.  pressure  in  a  wire-wound  tank.  When  used  as  a  starter 
it  drives  the  engine  through  a  train  of  spur  gears  and  rotates  at 
seven  times  engine  speed.  Automatic  arrangements  are  provided 
for  throwing  out  the  compressor  when  the  air  pressure  reaches 
230  Ib.  and  for  throwing  out  the  cranker  when  the  engine  starts. 
The  apparatus  weighs  about  50  Ib.  for  a  200-h.p.  engine. 


FIG.  338. — Compressed  air  distributor. 

In  the  Christensen  system  compressed  air  is  sent  through  a 
special  carburetor,  a  distributor  and  non-return  valves  to  the 
cylinders.  The  air  pressure  is  sufficient  to  start  the  engine 
(say  100  Ib.)  and  a  retarded  spark  gives  a  late  explosion  and 
initiates  regular  operation.  This  system  uses  a  minimum  of 
compressed  air  for  the  starting.  An  air  compressor  driven  by  a 
hand-operated  clutch  from  the  crankshaft  and  an  air  tank  are 
necessary  parts  of  the  system.  The  whole  weighs  about  40  Ib. 
for  a  six-cylinder  engine. 


432 


THE  AIRPLANE  ENGINE 


Compressed  air  can  be  used  inside  the  cylinders  without 
carbureting.  If  no  air  compressor  is  provided  a  steel  air  tank 
must  be  carried  of  sufficient  capacity  for  several  starts.  This 
arrangement  uses  a  distributor  and  non-return  valves;  it  will 
weigh  less  than  the  Christensen  system,  but  is  likely  to  leave  the 
aviator  stranded  on  occasion. 

The  details  of  an  air  distributor  are  shown  in  Fig,  338.  When 
the  starting  lever,  H,  is  thrown  in,  the  central  main  air  valve, 
Bj  is  opened  by  the  boss  at  the  end  of  the  sliding  sleeve  and  the 
individual  valves,  D,  E,  admitting  air  to  the  individual  cylinders 
are  opened  in  turn  by  the  rotation  of  the  face  cam,  G. 

Another  method  which  has  been  used  but  is  now  abandoned 
is  to  insert  a  black  powder  cartridge  in  a  special  fitting  in  the 
cylinder  head.  On  detonation  this  will  give  a  considerable 
pressure  and  should  start  the  engine.  It  causes  a  deposit  of 
carbon  in  the  cylinder  and  is  not  adapted  to  remote  control. 

Electric  cranking  has  been  used  considerably.  It  requires  a 
12-volt  battery  which  will  normally  have  to  supply  100  amperes 
but  may  be  called  upon  to  give  to  200  amperes  for  a  half  minute 
or  more;  an  electric  motor  which  will  operate  at  3,000  to  4,000 

r.p.m.,  carrying  a  small 
spur  gear  on  the  armature 
shaft;  and  a  double  reduc- 
tion gear  with  a  speed 
reduction  ratio  of  100  to 
150.  The  last  gear  is  pref- 
erably faster  °  1  f o  Xl 
propeller  hub 
motor  reduction  gears 
being  mounted  on  the 
crankcase  between  the 


FIG.  339. — Electric  starter. 


front  cylinders  and  the  propeller  hub.  The  gears  can  be  brought 
into  mesh  by  the  use  of  a  solenoid  wired  in  series  with  the  motor. 
The  Bijur  starter  used  on  Liberty-12  engines  has  a  six-cell  battery 
weighing  35  Ib.  with  a  capacity  of  starting  the  engine  150  times 
under  normal  conditions.  The  starting  motor  and  gear  may 
weigh  24  Ib.  and  give  an  engine  speed  of  40  to  50  r.p.m.  The 
maximum  torque  available  at  the  engine  crankshaft  is  1,300 
Ib.-ft.  An  arrangement  in  which  the  gears  are  kept  inside  the 
nose  of  the  engine  is  shown  in  Fig.  339  ;x  this  arrangement  sup- 
1  SHERMAN,  loc.  cit. 


STARTING  433 

poses  the  starter  incorporated  in  the  engine  design  and  not 
adapted,  as  usual,  as  an  afterthought.  If  this  starter  is  mounted 
at  the  rear  of  the  engine  it  can  be  made  a  combined  hand  and 
electric  starter  by  putting  a  dog  clutch  and  cranking  handle  on 
the  intermediate  shaft. 

Various  portable  crankers  have  been  devised  for  airdrome  use. 
The  U.  S.  Air  Service  has  used  an  electric  cranker  driven  by  an 
automobile-starting  motor  and  storage  battery  through  double- 
reduction  gearing,  and  mounted  on  a  motor  truck.  As  the  weight 
of  the  gearing  does  not  have  to  be  considered  in  this  case  it 
has  been  found  unnecessary  to  provide  an  automatic  release  in 
case  of  engine  back  fire.  The  cranker  is  mounted  in  a  spherical 
bowl  permitting  universal  adjustment;  it  is  brought  up  to  the 
end  of  the  propeller  shaft  and  adjusted  so  that  its  shaft  is  in  line 
with  the  engine  crankshaft.  An  engagement  lever  then  pushes 
the  perforated  face  plate  at  the  end  of  the  cranker  shaft  against 
the  front  propeller  hub  flange  when  some  of  the  nuts  enter  the 
perforations.  The  engagement  lever  is  withdrawn  as  soon  as 
the  engine  starts. 

The  Odier  portable  starter,  which  has  had  considerable  use, 
employs  a  long  single-acting  steel  cylinder  and  piston  operated 
by  carbon  dioxide  from  a  steel  bottle  of  liquefied  carbon  dioxide. 
The  piston  carries  a  pulley  at  its  free  end,  over  which  a  cable  is 
passed  with  one  end  fastened  to  the  cylinder  and  the  other 
wrapped  around  a  drum  and  then  fastened  to  an  elastic  cord. 
The  drum  has  four  bolts  placed  symmetrically  around  the 
periphery  at  one  end — parallel  to  the  drum  axis — projecting 
sufficiently  to  engage  a  kind  of  dog  clutch  on  the  front  propeller 
hub  flange.  The  cylinder  is  carried  on  an  inclined  wooden 
arm  and  a  vertical  leg  of  adjustable  height,  so  that  the  bolts  can 
be  brought  up  to  the  propeller  hub  level.  The  high  pressure  of 
saturated  vapor  of  carbon  dioxide  (308  Ib.  per  square  inch  at 
0°F.)  provides  a  large  starting  force  in  a  cylinder  of  small  diam- 
eter. The  piston  stroke  is  such  as  to  give  two  revolutions  of 
the  engine  with  high  speed.  The  cranker  is  thrown  forward  and 
out  of  mesh  when  the  engine  starts  by  the  action  of  the  dog  teeth 
on  the  propeller  flange.  In  case  of  a  back  fire  the  piston  is 
pulled  back,  the  gas  is  recompressed,  and  the  elastic  cord  becomes 
slack  and  permits  the  drum  to  revolve  freely.  The  weight  of  the 
whole  apparatus  is  44  Ib.  so  that  it  can  be  carried  in  the  plane  if 
desired. 

28 


CHAPTER  XIX 


POTENTIAL  DEVELOPMENTS 

Increased  Compression. — The  best  airplane  engines  give  a 
notably  better  performance,  both  as  regards  fuel  consumption 
and  horse  power  per  unit  of  piston  displacement,  than  other 
gasoline  engines.  The  possiblity  of  this  improved  performance 
results  from  the  use  of  a  higher  compression  ratio,  which  in  turn 
is  only  possible  through  the  use  of  the  volatile  aviation  gasoline. 
Ordinary  commercial  gasoline  containing  larger  fractions  of  the 
heavier  paraffines  (nonane  and  decane)  would  detonate  at  the 
compression  pressures  reached  in  airplane  engines.  The  de- 
pendence of  fuel  consumption  on  the  compression  ratio  is  shown 
in  the  following  table,1  which  gives  the  theoretical  consumption  of 
gasoline  (lower  heat  value  18,600  B.t.u.  per  pound)  per  brake 
horse-power  hour  with  an  assumed  mechanical  efficiency  of  90 
per  cent  and  with  variable  specific  heats. 


Ratio  of  compression  

4  0 

4  5 

5.0 

5.5 

6.0 

Gasoline  per  brake  horse-power 
hour  

0  416 

0  398 

0  375 

0.361 

0.350 

The  best  modern  engines  use  0.45  to  0.50  Ib.  per  brake  horse- 
power hour  with  a  compression  ratio  around  5.0;  that  is,  the 
attained  efficiency  is  0.375  -f-  0.45  =  83  per  cent  of  the  theore- 
tical efficiency.  With  higher  compressions  this  relative  efficiency 
tends  to  increase.  At  maximum  load,  which  is  obtained  only 
with  richer  mixtures  (see  p.  33),  the  relative  efficiency  averages 
75  per  cent  for  water-cooled  and  76.5  per  cent  for  radial  air-cooled 
engines  with  aluminum  cylinders. 

The  use  of  still  higher  compression  pressures,  and  consequently 
higher  efficiencies,  is  possible  by  the  use  of  gasoline  mixed  with 
toluol,  alcohol  or  other  substance  which  will  prevent  detonation 
(see  p.  236).  This  field  of  improvement  is  now  under  active 
investigation  and  promises  considerable  improvement. 

1  GIBSON,  Trans.  Royal  Aeronautical  Society,  1920. 

434 


POTENTIAL  DEVELOPMENTS 


435 


Use  of  Inert  Gases. — A  high  compression  pressure  can  be  used 
without  danger  of  detonations,  and  consequent  preignitions,  by 
taking  in  cooled  exhaust  gases  with  the  charge.  The  influence 
of  such  admixture  is  shown  in  Fig.  340,  which  is  taken  from 
Ricardo's  tests.1  With  a  high-grade  fuel  which,  when  operating 
at  full  throttle  and  with  an  economical  setting,  detonates  at  a 
compression  ratio  of  4.85:1,  the  full  power  can  be  maintained  by 
admitting  inert  gases  in  sufficient  quantity  to  prevent  detonation 
up  to  a  ratio  of  compression  of  6:1.  That  is,  the  decrease  in 
weight  of  fresh  charge  taken  in  is  fully  compensated  by  the  in- 
crease in  engine  efficiency  up  to  that  ratio  of  compression.  If 
still  higher  compression  is  used  (for  an  oversized  engine  for  high- 


1 


150 
140 
130 


g 

£  0.6 


0.4 


0.3 


40      4.5       5.0       5.5      60      6.5      7.0       7.5     8.0 
Compression  Ratio. 

FIG.  340. — Effect  of  addition  of  inert  gases  on  engine  performance. 

altitude  flight,  see  p.  390)  the  power  will  fall  off.  The  increase 
in  economy  by  the  use  of  the  inert  gases,  with  increase  of  com- 
pression ratio  from  4.85  to  6,  is  seen  to  be  about  6  per  cent.  The 
dotted  lines  show  the  performance  obtained  using  a  fuel  doped  to 
prevent  detonation  and  without  the  admission  of  inert  gases. 
With  a  compression  ratio  of  7:1  the  exhaust-controlled  engine 
develops  84  per  cent  of  the  power  which  would  be  developed  by  a 
pure  non-detonating  mixture. 

Any  increase  in  efficiency  from  increase  of  compression  ratio 
will  also  increase  the  power  output  in  practically  the  same  ratio 
and  is  therefore  doubly  valuable. 

1  Trans.  Royal  Aeronautical  Society,  1920. 


436  THE  AIRPLANE  ENGINE 

The  maximum  brake  m.e.p.  possible  for  an  engine  with  inlet 
valve  closure  of  50  deg.  late,  with  volumetric  efficiency  of  88  per 
cent  and  mechanical  efficiency  of  90  per  cent,  is  as  follows:1 


Nominal  compression 

ratio       

4 

5 

5. 

0 

5 

5 

6 

0 

Maximum  brake  m.e. 

P  

130 

5 

138. 

0 

143 

5 

148 

0 

The  best  recorded  results  for  both  air-  and  water-cooled  engines 
with  compression  ratios  of  4.5  to  5.0  are  very  close  to  these 
figures. 

Tests  of  single  engines  have  shown  consistently  better  results 
than  those  of  multi-cylinder  engines.  The  best  air-cooled  single- 
cylinder  engines  have  shown  a  relative  thermal  efficiency  of  91 
per  cent.  The  difference  between  the  best  performance  of  a 
single-cylinder  water-cooled  engine  and  the  performance  of  a 
12-cylinder  Vee  engine  is  from  8  to  10  per  cent.  The  difference 
depends  on  the  efficiency  of  the  induction  system  and  represents 
the  possible  saving  by  better  distribution  in  the  induction  system. 

As  the  efficiencies  of  the  best  single-cylinder  engines  using 
weak  mixtures  are  within  10  to  15  per  cent  of  the  theoretical 
maximum,  it  is  evident  that  little  further  progress  is  possible  in 
improving  the  thermal  efficiency  of  engines  using  the  present 
cycle  of  operations.  Increase  in  capacity  (b.h.p.  per  unit  of 
piston  displacement)  can  be  obtained  by  supercharging  or  by  im- 
proving the  volumetric  efficiency.  This  latter  method  offers  some 
chance  for  improvement  as  the  measured  values  vary  from  70  to 
85  per  cent  with  exceptional  values  up  to  90  per  cent. 

The  fire  risk  in  airplanes  could  be  practically  eliminated  by 
the  use  of  kerosene  as  fuel.  Kerosene  may  be  used  either  (1)  by 
vaporizing  it  outside  the  engine,  or  (2)  injecting  it  into  the  cylin- 
der as  a  liquid  either  during  the  suction  stroke  or  at  the  end  of 
compression.  To  vaporize  a  reasonable  proportion  the  initial 
temperature  must  be  not  less  than  140°F.,  which  results  in  a 
reduction  in  the  weight  of  the  charge  of  about  20  per  cent  as 
compared  with  gasoline  and  a  corresponding  decrease  in  engine 
power.  Furthermore,  it  is  chemically  much  less  stable  than  gaso- 
line and  detonates  at  a  lower  temperature  so  that  a  lower  com- 
pression ratio  must  be  used,  which  further  diminishes  the  power 
and  decreases  the  efficiency.  The  heavier  fractions  condense 
on  the  cylinder  wall  and,  passing  into  the  crankcase,  thin  the 

1  GIBSON,  loc.  tit. 


POTENTIAL  DEVELOPMENTS  437 

lubricating  oil.  Injection  of  the  fuel  during  the  suction  stroke 
intensifies  this  last  trouble  but  reduces  the  loss  due  to  preheating. 
Injection  at  the  end  of  compression  presents  many  difficulties 
common  to  Diesel  engines,  which  are  not  yet  satisfactorily  solved. 

Modifications  of  the  Otto  cycle  would  seem  to  offer  considerable 
possibilities  for  increased  efficiency.  The  pressure  at  the  opening 
of  the  exhaust  valve  is  usually  from  60  to  70  Ib.  per  square  inch. 
If  more  complete  expansion  could  be  obtained  a  considerable 
increase  in  efficiency  might  be  effected.  Attempts  have  been 
made  to  realize  this  potential  increase  in  work  along  two  lines,  (1) 
by  more  complete  expansion  in  the  cylinder  and  (2)  by  expanding 
the  gases  after  leaving  the  cylinder. 

1.  A  lower  terminal  pressure  in  the  cylinder  can  be  obtained 
either  (a)  by  throttling  or  cutting  off  the  admission  of  the  charge 
or  (b)  by  making  the  expansion  stroke  longer  than  the  com- 
pression stroke.  The  former  method  (a)  is  that  of  the  oversized 
engine  (p.  390)  and  is  employed  primarily  for  maintaining  power 
at  high  altitudes.  Its  use  requires  a  larger  and  therefore  heavier 
engine  for  a  given  capacity,  which  is  a  serious  detriment.  The 
method  (b)  can  be  carried  out  by  the  use  of  a  variable -stroke 
engine  and  has  the  incidental  advantage  of  permitting  a  more 
complete  scavenging  of  burned  gases. 

The  Zeitlin  engine  which  is  now  under  development  is  an 
example  of  a  variable  stroke  engine,  but  this  feature  is  utilized  in 
this  case  only  to  give  better  scavenging  and  thereby  to  permit  the 
admission  of  a  greater  weight  of  charge.  It  is  a  single-valve, 
nine-cylinder  air-cooled  rotary  which  follows  the  Gnome  engine  in 
taking  in  its  air  supply  through  the  open  exhaust  valve  and 
mixing  with  it  an  overrich  mixture  through  ports  uncovered  by 
the  piston  near  the  end  of  the  suction  stroke.  The  variable 
stroke  is  obtained  by  mounting,  on  the  crankpin,  eccentrics,  which 
are  driven  by  gearing  around  the  crankpin  in  the  same  direction 
as  the  engine  but  at  one-half  engine  speed.  The  connecting  rods 
are  mounted  on  these  eccentrics  and  consequently  the  piston 
travel  will  vary  throughout  two  revolutions  of  the  engine.  In  the 
engine  with  107.75  mm.  crank  throw,  the  working  stroke  of  the 
engine  is  181  mm. ;  the  exhaust  or  scavenging  stroke  is  203.5  mm.  ; 
the  suction  stroke  is  226  mm.  and  the  compression  stroke  203.5 
mm.  As  the  admission  ports  are  not  covered  until  the  piston  has 
made  part  of  the  compression  stroke  the  effective  compression 
stroke  is  practically  the  same  as  the  working  stroke.  The  admis- 


438  THE  AIRPLANE  ENGINE 

sion  ports  in  the  cylinder  are  uncovered  near  the  end  of  the  suc- 
tion stroke.  The  length  of  the  scavenging  stroke  is  such  as  to 
clear  the  cylinder  almost  completely  of  burned  gases,  so  that  the 
new  charge  is  undiluted  by  them. 

This  engine  is  arranged  to  give  variable  compression  by  open- 
ing the  exhaust  valve  during  the  early  portion  of  the  compression 
stroke  and  permitting  some  air  to  escape  before  it  has  time  to 
mix  with  the  overrich  mixture.  It  is  designed  for  a  maximum 
compression  ratio  of  7  and  has  controls  for  decreasing  this  to 
4.5.  The  maximum  compression  is  for  altitudes  of  10,000  ft.  or 
more;  the  minimum  compression  is  for  ground  operation.  The 
strength  of  the  mixture  will  obviously  change  with  the  ratio  of 
compression;  the  mixture  will  be  rich  at  the  ground  and  will  be 
leaned  to  maximum  economy  strength  at  normal  flying  level. 

2.  The  further  expansion  of  the  gases  after  leaving  the  cylinder 
can  be  carried  out  either  in  a  gas  turbine  or  in  a  reciprocating 
engine.  The  use  of  the  exhaust  gas  turbine  for  driving  a  super- 
charging blower  has  already  been  discussed  (p.  408) ;  in  view  of 
the  high  speed  of  the  turbine  shaft,  which  is  necessary  if  the 
turbine  is  to  have  fair  efficiency,  it  is  doubtful  whether  such  a 
turbine  could  be  geared  down  so  as  to  help  drive  the  propeller 
without  excessive  loss  of  power.  For  driving  high-speed  auxil- 
iaries such  as  the  supercharging  blower  it  has  an  important 
field. 

Compounding. — Expansion  of  the  gases  in  a  reciprocating 
engine  can  be  accomplished  by  following  the  methods  used  in 
compound  steam  engines.  The  problem  is  simplified  in  some 
respects  in  the  gasoline  engine  because  one  double-acting  low- 
pressure  cylinder  can  take  the  exhaust  from  four  high-pressure 
cylinders  and  as  the  temperature  of  the  gases  is  greatly  reduced 
by  the  time  they  reach  the  low-pressure  cylinder  it  might  be 
possible  to  operate  without  trouble  from  excessive  piston  tem- 
perature. As  the  friction  work  in  the  high-pressure  cylinders  is 
equal  to  about  15  Ib.  per  square  inch  of  piston  area  there  should 
be  a  pressure  drop  of  at  least  that  amount  at  exhaust,  or  with 
70  Ib.  terminal  pressure  in  the  high-pressure  cylinder  the  receiver 
pressure  should  be  about  50  Ib.  Furthermore,  in  order  to  get  a 
full  charge  in  the  high-pressure  cylinder  it  would  be  necessary 
to  have  an  atmospheric  exhaust  from  the  high-pressure  cylinder 
at  the  end  of  the  exhaust  stroke  and  immediately  after  closure 
of  the  exhaust  to  the  receiver.  The  piston  displacement  of  the 


POTENTIAL  DEVELOPMENTS  439 

low-pressure  cylinder  would  probably  have  to  be  about  three 
times  that  of  the  high-pressure  cylinder,  or  with  the  same  stroke 
its  diameter  would  be  1.7  times  that  of  the  high-pressure  cylin- 
der. Attempts  to  construct  compound  gas  engines  have  been 
made  in  stationary  types  but  without  commercial  success.  The 
extra  power  obtainable  has  not  justified  the  additional  first  cost 
and  maintenance.  It  is  possible,  however,  that  more  success 
may  be  met  in  the  airplane  engine  where  first  cost  is  not  of  prime 
importance.  The  weight  of  the  engine  per  horse  power  should 
not  be  increased  by  compounding  and  the  weight  of  fuel  used 
should  be  appreciably  diminished. 

Two-cycle. — A  modification  of  the  Otto  cycle  which  offers 
possibilities  of  considerable  reduction  in  weight  per  horse  power 
is  two-cycle  operation.  The  normal  Otto  cycle  requires  four 
strokes  for  the  completion  of  the  cycle  of  which  two  are  used 
solely  for  pushing  out  the  burned  charge  and  taking  in  the  new 
charge.  The  essential  parts  of  the  cycle  are  unchanged  if  the 
exhaust  and  admission  processes  are  speeded  up  and  made, 
in  part,  simultaneous  by  using  a  slightly  precompressed 
charge  to  sweep  out  the  exhaust  gases  after  the  exhaust  pressure 
has  fallen  substantially  to  atmospheric  pressure.  With  this 
arrangement  the  cycle  of  operation  may  be  completed  in  two 
strokes,  the  number  of  cycles  per  minute  doubled  and  the  horse 
power  almost  doubled.  Considerable  experience  with  two- 
cycle  engines  is  available  from  stationary  practice  and  marine 
practice  and  indicates  the  possibility  of  increase  of  the  power 
output  from  a  cylinder  of  given  size  by  from  60  to  80  per  cent  but 
with  a  falling  off  in  efficiency  of  20  per  cent  or  more. 

There  are  two  general  methods  of  obtaining  a  precompressed 
charge,  (1)  by  crankcase  compression  and  (2)  by  the  use  of  a 
separate  compressor.  With  crankcase  compression  in  a  multi- 
cylinder  engine,  the  crankcase  must  be  divided  to  form  a 
gas-tight  compartment  for  each  cylinder  so  that  each  piston  on 
its  down  stroke  may  compress  a  charge  which  has  been  taken  in 
during  the  up  stroke.  This  arrangement  is  only  possible  in  an 
engine  with  a  single  row  of  cylinders;  it  cannot  be  carried  out  in 
a  Vee,  W,  or  radial  engine.  It  is  simpler  than  the  separate 
compressor  but  it  limits  the  amount  of  charge  which  can  be 
taken  in  to  the  volume  sucked  in  during  the  up  stroke  of  the 
piston  and  it  will  carry  lubricating  oil  from  the  crankcase  into 
the  cylinder.  A  separate  compressor  should  have  a  displacement 


440 


THE  AIRPLANE  ENGINE 


volume  greater  than  that  of  the  engine  cylinder  in  order  to  send 
some  scavenging  air  into  the  engine  cylinder  for  the  more  com- 
plete clearing  out  of  the  exhaust  gases  and  also  to  give  a  pressure 
at  the  beginning  of  compression  which  is  fully  up  to  or  slightly 
above  atmospheric  pressure.  The  compression  pressure  required 
is  about  5  Ib.  per  square  inch.  One  double-acting  air  compressor 
would  be  required  for  two  engine  cylinders  with  discharge  from 
the  compressor  direct  to  the  cylinders.  By  the  use  of  a  receiver 
a  larger  compressor  can  be  made  to  serve  a  larger  number  of 
cylinders.  A  centrifugal  compressor  would  eliminate  the  need 
for  a  receiver. 

The  two-cycle  engine  is  just  beginning  to  be  used  in  airplane 
service.  The  principal  difficulties  to  overcome  are  low  efficiency 
and  heat  trouble.  If  the  additional  weight  of  fuel  that  has  to  be 


Annular  Exhaust 


&  Air Inkt  Ports- 
Exhaust  Pork 


tool  ing  Water  Inlet 


FIG.  341. — Junkers  two-cycle  solid-injection  engine. 

carried  for  a  long  flight  is  equal  to  the  saving  in  engine  weight 
there  is  little  advantage  in  the  lighter  weight  engine  except  for 
short  flights.  The  doubled  number  of  explosions  in  the  engine 
per  unit  of  time  increases  the  cylinder  temperature  and  leads  to 
serious  heat  difficulties  with  the  piston.  Exhaust  valve  troubles 
are  eliminated  by  the  use  of  exhaust  ports  uncovered  by  the 
piston  in  place  of  exhaust  valves.  There  has  been  considerable 
experimental  work  in  this  field  but  with  no  practical  results  so  far 
except  in  the  case  of  the  Junkers  engine. 

The  Junkers  engine  (Fig.  341)  obtains  high  efficiency  and 
eliminates  heat  trouble  by  departing  entirely  from  the  usual 
construction.  There  are  six  horizontal  cylinders  with  their 
axes  at  right  angles  to  the  center  line  of  the  fuselage.  The 
engine  has  two  opposed  pistons  per  cylinder  and  two  crankshafts. 
All  the  pistons  on  each  side  of  the  engine  are  connected  to  a 


POTENTIAL  DEVELOPMENTS  441 

common  crankshaft.  The  two  crankshafts  are  geared  together 
so  as  to  make  the  two  pistons  of  any  one  cylinder  move  in  or  out 
simultaneously.  The  combustion  chamber  is  the  space  enclosed 
between  the  two  pistons  when  they  are  on  their  inner  dead 
center.  Near  the  outer  dead  centers  the  pistons  uncover  cylinder 
ports,  the  exhaust  ports  (on  the  left)  being  uncovered  first  and 
the  air  admission  ports  (on  the  right)  shortly  afterwards.  The 
propeller  shaft  carries  the  central  gear  with  which  the  gears  on 
the  two  crankshafts  mesh;  a  blower  is  operated  from  the  pro- 
peller shaft.  There  are  no  valves  and  no  carburetors.  When 
the  pistons  are  on  their  outer  dead  centers,  air  from  the  blower 
passes  through  the  cylinder  from  right  to  left  and  clears  out  the 
exhaust  gases.  This  air  is  compressed,  while  the  pistons  make 
their  inward  strokes,  to  a  pressure  of  210  Ib.  per  square  inch  or 
more.  Fuel  is  then  injected  into  the  combustion  space 
through  the  nozzle  at  the  bottom  of  the  cylinder,  is  ignited  by  a 
spark  plug  immediately  above  it,  and  expands,  driving  the  two 
pistons  outward.  The  pistons  are  equipped  with  a  special 
cooling  device.  They  are  made  with  a  cavity  which  is  partly 
filled  with  a  heavy  oil  and  then  sealed.  The  oil  is  violently 
dashed  backwards  and  forwards  by  the  motion  of  the  piston, 
absorbs  heat  from  the  piston  head  and  carries  it  to  the  cooled 
piston  sides.  The  efficiency  of  this  method  of  cooling  is  shown 
by  tests  with  thermo-elements  which  indicated  a  maximum 
temperature  of  the  piston  head  of  350°F.  at  maximum  speeds 
and  loads. 

The  advantages  of  this  method  of  construction  appear  to  be 
manifold.  The  high  compression  gives  high  efficiency,  which  is 
helped  by  the  small  heat  loss  during  explosion  resulting  from  the 
smallness  of  the  cooling  surface  of  the  combustion  chamber.  The 
excellent  scavenging  permits  higher  volumetric  efficiency,  which 
in  conjunction  with  higher  efficiency  gives  a  higher  m.e.p.  than 
is  obtainable  in  other  types.  The  i.h.p.  is  consequently  more 
than  twice  that  obtained  for  the  same  piston  displacement  in 
four-cycle  engines  and  the  weight  is  reduced  to  1.5  Ib.  per  horse 
power.  The  balancing  of  the  reciprocating  parts  is  practically 
perfect  because  the  two  pistons  of  each  pair  are  at  all  times 
moving  with  equal  accelerations  in  opposite  directions;  this 
condition  is  favorable  to  high  engine  speeds.  The  large  size  of 
the  gas  inlet  and  exhaust  ports  combined  with  positive  admission' 
of  the  air  and  fuel  permits  also  high  volumetric  efficiency  at 


442  THE  AIRPLANE  ENGINE 

high  speeds;  consequently  this  engine  can  be  operated  at  higher 
speeds  than  the  usual  type.  As  the  propeller  is  geared  it  can  be 
run  at  its  most  favorable  speed.  A  further  feature  of  the  engine 
is  the  great  reduction  of  fire  risk  resulting  from  the  direct  dis- 
charge of  the  fuel  into  the  engine  cylinder;  there  is  no  explosive 
mixture  outside  the  engine  cylinder  and  less  liability  to  fuel 
leaks.  Moreover,  a  less  volatile  fuel  can  be  burned. 

The  principal  apparent  objection  to  the  Junkers  engine  is  the 
difficulty  of  accommodating  an  engine  of  its  width  in  the  fuselage. 
In  the  design  shown  in  Fig,  341,  the  over-all  width  is  8J-£  times  the 
stroke,  or  with  a  6-in.  stroke  the  width  is  4  ft.  3  in.  This  dimen- 
sion is  exceeded  in  large  radial  engines  (see  p.  195).  The  mechan- 
ical efficiency  is  low  on  account  of  the  blower  work  and  of  the 
gearing  losses;  other  Junkers  engines,  not  adapted  to  airplane 
use,  have  given  mechanical  efficiencies  of  about  73  per  cent.1 

Among  the  possible  developments  for  airplane  engines  are 
Diesel  engines,  gas  turbines  and  steam  plants  either  turbine 
or  reciprocating.  The  Diesel  engine  would  be  most  advantage- 
ous in  view  of  its  higher  efficiency  and  the  safety  and  low  cost  of 
the  fuels  which  it  could  employ.  The  difficulties  in  the  way  of  its 
employment  in  airplanes  in  its  present  stage  of  development  are 
its  excessive  weight  and  the  large  size  of  individual  cylinder 
below  which  it  has  not  been  found  practicable  to  go.  In  the 
modified  form  of  the  Hvid  and  similar  engines,  smaller  size 
cylinders  become  practicable  but  the  weight  is  still  excessive. 
It  is  by  no  means  certain  that  it  will  be  found  practicable  to 
operate  this  cycle  successfully  at  the  high  speeds  necessary  for 
airplane  use. 

Gas  turbines  have  been  under  active  development  for  over 
fifteen  years  but  the  difficulties  inherent  in  them  have  not  as  yet 
been  overcome  without  sacrificing  their  potential  efficiencies. 
The  principal  troubles  are  those  resulting  from  the  high  temper- 
atures to  which  the  combustion  chamber,  nozzles,  and  buckets 
are  subjected.  When  these  temperatures  are  reduced,  by  inject- 
ing water  or  excess  air,  the  efficiencies  fall  off.  Moreover,  the 
efficiencies  are  low  unless  the  air  is  precompressed  and  as  centrif- 
ugal compressors  seldom  have  efficiencies  above  about  60  per 
cent,  a  large  part  of  the  power  developed  in  the  turbine  is  utilized 
in  driving  the  compressor.  Over-all  thermal  efficiencies  are 
usually  about  5  per  cent,  although  an  unsubstantiated  value  of 

1  SCOTT,  Jour.  Soc.  AuL  Eng.,  1917. 


POTENTIAL  DEVELOPMENTS  443 

20  per  cent  has  been  claimed  for  a  1,000-h.p.  unit.  The  gas 
turbine,  if  applied  to  airplane  propulsion,  would  have  to  be  geared 
down,  probably  with  double  reduction  gear.  Its  simplicity  and 
light  weight  have  attracted  many  inventors,  but  there  are  no 
indications  that  it  is  ever  likely  to  become  practically  available. 
Steam-power  plants  have  the  great  advantage  over  internal 
combustion  engines  of  maintaining  their  power  at  all  altitudes. 
Both  turbines  and  reciprocating  engines  of  light  weight  are 
fully  developed  and  can  be  regarded  as  immediately  available 
for  airplane  use.  The  difficulties  arise  in  connection  with  the 
boiler  and  oil  burner.  For  efficiency  the  steam  must  be  generated 
at  high  pressure  and  with  high  superheat,  but  no  construction  of 
boiler  is  known  which  does  not  entail  weights  which  would  be 
excessive  for  aircraft.  A  satisfactory  kerosene  (or  other  fuel) 
burner  for  operation  with  high  rates  of  combustion  in  small 
space  would  also  have  to  be  developed;  the  fire  risk  from  such 
apparatus  would  probably  be  considerable.  And  finally,  the 
efficiency  of  the  best  steam  plants  is  not  nearly  so  high  as  that  of 
existing  airplane  engines.  It  seems  very  unlikely  that  steam 
plants  will  ever  be  employed  in  aircraft. 


INDEX 


A.  B.  C.  engines,  (table),  194 
Acceleration,  in  carburetors,  271 

force,  (see  Inertia  Force} 
Acetylene,  as  fuel,  properties  of,  225 
Aerofoil,  (see  Wings) 
Air  compression,  power  absorbed  in, 

397 

temperature  rise  in,  395,  400 
cooling,  344 
cycle  efficiency,  (def.  and  table), 

17 
densities,    at    altitudes,    (curves), 

366 

flow,  pulsating,  254 
'flow  of,  orifice  coefficients  for,  243 
through  venturi   tube,    246- 

250 
fuel  ratio,   determination  of 

strength  of,  242 
influence   of  charge  dilution 
on  optimum  value  of,  262 
influence  of,  on  engine  per- 
formance, 260 
optimum  values  of,  261 
variation  with  air  density  of, 

259 

intakes,  416 
pumps,  292 

starting,  (see  also  Starting'),  429 
temperature,  at  altitude,  (curves), 

365 

influence  of,  on  capacity,  35 
on  engine  power,  33,  35 
on  thermal  efficiency,  35 
weight  flow  of,   (chart),  249 
Alcogas,  242 

Alcohol,  as  fuel,  properties  of,  224 
effect  on  detonating  pressure,  239 
mixtures  with  gasoline,  242 
Alloys    (see   Steels,    and   Aluminum 
Alloys) 


Altitude,  air  density  at,  (curves),  366 
control  of  carburetor,  268 
effect  of,  on  radiator,  365 
influence    of,    on    engine    power, 

386-389 

temperatures  at,   (curves),  365 
Aluminum  alloys,  118 

for  pistons,  135 
American    engines,    description   of, 

80-98 
Austro -Daimler  engine,  cylinder  of, 

131 

dimensions  of,  73,  124 
inertia  forces  and  bearing  loads, 

59 

water  pumps  of,  372 
weights  of  parts,  78 


B 


B.   R.  2  engine,  description  of,  190 
Balancing,  devices  for,  58 
in  radial  engines,  206 
in  rotary  engines,  206 
of  reciprocating  parts,  55 
of  rotating  parts,  54 
Ball  and  ball  carburetor,  description 

of,  283 

Battery  ignition  systems,  (see  Igni- 
tion Systems) 

Battery,  of  Liberty  engine,  316 
Basse-Selve  carburetor,  description 

of,  284 

engine,  compression  release  of,  428 
cylinder  dimensions  of,  124 
dimensions  of,  73 
inertia  forces  and  bearing  loads 

of,  59 

lubricating  system  of,  341 
oil  pumps  of,  341 
valve  gear  of,  171 
valves  of,  157    . 
weights  of  parts,  78 


445 


446 


INDEX 


Baume"  scale,  conversion  table  for, 

219 

Bayerische  Motoren  Werke  carbu- 
retor,   description   of,    285 
Bearings,  144 

ball  and  roller,  in  radial  engines, 

207 

crankshaft,  dimensions  of,  74 
friction  work  of,  328 
loads  on,  59,  76,  327 
oil  grooving  of,  337 
Benz  engine,  compression  release  of, 

427 

connecting  rods  of,  130,  141 
cylinder  of,  124,  129 
description  of,  110 
dimensions  of,  73 
fuel  pump  of.  294 
intake  manifold  of,  420 
lubrication  system  of,  335 
oil  pump  of,  336 
performance  curves  of,  111 
piston  of,  136 
propeller  hub  of,  148 
valve  gear  of,  172 
weights  of  parts,  78 
Benzene,  (see  Benzol) 
Benzol,  as  fuel,  properties  of,  223 
Bijur  starting  system,  432 
Bosch  magneto,  302 
Brake  mean  effective  pressure,  (def.), 

25 

Bugatti  engine,  description  of,  93-98 
performance  curves  of,  98 
spark  adjustment  in,  306 
water  pumps  of,  372 

C 

Cams,  163 

followers  for,  164 
Camshafts,  dimensions  of,  72 
Capacity  of  engine,  26,  29 
influence  of  fuels  on,  240 
variation   with   air   temperature, 

33,  35 

compression  ratio,  37,  392 
engine  speed,  37,  392 
jacket-water  temperature,  37 
mixture  strength,  33 


Carbon  dioxide,  dissociation  of,  20 
Carburetors,  245-289 
acceleration  in,  271 
air  discharge  coefficients  of,  252 
altimetric  compensation  of,  267 

control  of,  268 
atomization  in,  271 
construction  of,  272 
dimensions  of,  74 
float  arrangements  in,  272 
flooding  of,  272 
idling  device  in,  271 
intakes  for,  272 
mixture  characteristics  of,  258 
performance  of,  264 
pressure  drop  in,  253 
strainers  for,  290 
viscous  flow  type,  267 
weights  of,  78 
Castor  oil,  333 
Central  power  plants,  384 
Choke,  (see  also  Venturi  Tube),  250 
Christensen  system  of  starting,  431 
Claudel  carburetor,   description  of, 

276 

Clerget  engine,  description  of,  189 
effect  of  compression  ratio  on 

horsepower  of,  393 
performance  curves  of,  190 
Combustion,  air  required  for,  228 
effect  of  turbulence  on  rate  of,  235 
higher  and  lower  heats  of,  (def.), 

214 

products  of,  (table),  216 
velocity  of  propagation  of,  232 
Compound  engines,  438 
Compression,  ratio  of,  (def.),  14 

variable,  438 

pressures,  effect  of  alcohol  on,  239 
effect  of  toluene  on,  239 
maximum  allowable,  236,  237 
variation      with      compression 

ratio,  16 
engine  speed,  16 
ratio,  66,  72 

influence  of,  on  capacity,  37 
on  efficiency,  434 

engine      power      at      high 
altrtudes,  392 


INDEX 


447 


Compression,  release,  427 
Compressors,  (see  Air  Compression} 
Connecting  rod,  rods,  138 

assembly,  articulated,  203 

dimensions  of,  74 

for  rotary  and  radial  engines, 

203 

materials  for,  116 
slipper  assembly,  204 
stresses  in  articulated,  143 
weights  of,  76,  78 
Cooling  fins,  345 
systems,  344 

anti-freeze  solutions  for,  375 
piping  for,  375 

pumps  for,   (see  Water  Pumps) 
typical  examples  of,  377 
water  for,  375 

Cosmos  engines,  (table),  195 
induction  chamber  of,  196 
"Jupiter,"  balancing  of,  207 
performance  curves  of,  196 
Counterweights,  146 
Crankcases,  148 
cooling  of,  150 
weights  of,  (table),  78 
Crank  pins,  loads  on,  76 
Crankshafts,  143-146 
balancing  of,  143 
dimensions  of,  74 
materials  for,  115 
of  radial  engines,  211 
strength  of,  146 
weight  of,  78 

Curtiss  engines,  crankshaft  balanc- 
ing of,  144 
cylinders  of,  124,  127 
description  of,  86-90 
lubricating  systems  of,  334 
performance  curves  of,  90,  92 
type  K,  86 

Cylinders,  air-cooled,  200,  345 
materials  for,  349 
temperature  of,  347] 
arrangements,    size    and    propor- 
tions of,  61 

attachment  to  crank  case,  125 
liners,  material  of,  115 
lubrication  of,    (see  Lubrication) 


Cylinders,  offset  of,  48 
thickness  of,  123 
types  of  construction,  122     g 
weights  of,  78 

D 

Detonation,  236 
Diesel  engines,  442 
Dilution    of    charge,    influence    of 
compression    pressure    on, 
263 

influence    on    engine    perform- 
ance, 393,  434 

Dimensions,  engine,  (table),  65 
of  American  and  German  engines, 

(table),  72 

overall,  of  engines,  (table),  79 
Dissociation,  of  carbon  dioxide,  20 

of  water  vapor,  20 
Distillation  curves,  241 
Distributor,  Bosch,  304 

Dixie,  305 
Dixie  magneto,  302 
Double-rotary  engines,  176,  191 
Drag,  of  wing,  2 
Duralumin,  121 


E 


Engine     speed,     influence     of,     on 

capacity,  37 
English    engines,     descriptions    of, 

98-107 

Ether,  as  fuel,  properties  of,  226 
Exhaust  gas,  composition  of,  244 
turbines,    (see  also  Supercharg- 
ing), 398,  438 
manifolds,  (see  Manifolds) 
mufflers,  (see  Mufflers) 
Explosion     phenomena,     (see     also 

Fuels,  Combustion),  231 
intervals,  in  multicylinder  engines, 

63 
limits,    effect   of   carbon   dioxide 

dilution  on,  262 
Explosive    mixtures,  212 

properties  of,  (table),  216 
wave,  236 


448 


INDEX 


Fiat  engine,  description  of,  107 
lubrication  system  of,  335 
valves  of,  174 
Firing  order,  75,  325 
Flash  point,  of  oils,  330 
Flight,  power  available  for,  8 

power  required  for,  1 
Friction,  laws  of,  327 
loss,  at  piston,  24 

in  engine,  24 
Fuel,  fuels,  212-226 

air  ratio,  (see  Air-fuel  Ratio) 

air  required  for  combustion   of, 

(table),  216 
detonating  compression  pressures 

of,  (table),  237 

explosive  limits  of  air-fuel  mix- 
tures, (table),  232 
flow  through  jets,  254-258 
heats  of  combustion  of,    (table), 

217 

ignition  temperatures  of,  233 
influence  on  capacity  of,  240 
influence  of  temperature  on 

fluidity,  (curves),  257 
hydrocarbon,  classification  of,  213 
minimum  vaporization  tempera- 
ture of,  231 
properties  of,  212,  215 
pumps,  289,  292 
weights  of,  78 
specific  heats  of,  229 
specific     volumes     of     saturated 

vapors  of,  229 
systems,  289-294 

for  supercharging  engine,  412 
tanks,  290 

temperature  drop  due  to  vapori- 
zation of,  230 
vapor  pressures  of,  229 
viscosity  of,  257 


G 


Gas  turbines,   (see  also  Exhaust  Gas 

Turbines),  442 
velocity,  through  valves,  152 


Gasoline,  (see  also  Fuel),  217-223 
blended  casing-head,  218 
calorific  value  of,  223,  241 
cracked,  (synthetic),  218 
distillation  curves  of,  221 
mixtures  with  alcohol,  242 
specifications  for,  219 
"straight"  refinery,  218 
tests  for,  222 
volatility  of,  220 

Geared  propeller  drives,  378-384 

Gears,  heating  of,  384 
pressure  between  teeth  of,  383 
stresses  in,  383 

Gnome  engine,  description  of,  185 
torque  of,  53 

Gudgeon  pin,  (see  Piston  Pin) 


Hall-Scott  engine,  cylinder  dimen- 
sions of,  124 
description  of,  90 
exhaust  manifold  of,  420 
lubrication  system  of,  335 
performance  curves  of,  95 
Heat  balance,  of   airplane   engine, 

38 

dissipation,  from  air-cooled  cylin- 
ders, 345 

transfer,  in  radiator  core,  352 
Hispano-Suiza  engine,  air  pump  of, 

292 

cylinder  of,  126 
description  of,  82 
dimensions  of,  73 
effect  of  compression  ratio  on 

horsepower  of,  393 
hand  starting  mechanism  of,  430 
influence    of    air    density    on 

performance,  388 
lubrication  system  of,  334 
performance  curves  of,  86 
propeller  gears  of,  378  , 
propeller  hub  of,  147 
starting  torque  of,  427 
test  results  of,  31 
valves  of,  173 
weights  of  parts  of,  78 


INDEX 


449 


Horse  power  (see  also  Capacity) 

required  for  flight,  4 
Hydrogen,  as  fuel,  properties  of,  225 


Idling  device,  in  carburetors,  271 
Ignition,  295-326 

assemblies,  weights  of,  78 
spark  advance,  306 
systems,  battery,  297,  312 
comparison  of,  324 
cycle  of  operations  in,  307 
self-sustaining  battery,  313 
temperature  of,  233 
Indicated  thermal  efficiency,  maxi- 
mum obtainable,  237 
Indicator  cards,  actual,  15 

negative  pumping  loop,  23 
theoretical,  13 
Induction  coil,  296 
Inertia  factors,  (table),  42 

forces,  in  Austro-Daimler,  Basse"- 
Selve  and  Liberty  engines, 
59 

in  radial  engines,  52 
primary,  55 
secondary,  55 
of  reciprocating  parts,  41 
Intake  manifolds,  (see  Manifolds) 


Jacket-water  temperature,  influence 

of,  on  capacity,  37 
Jets,  discharge  coefficients  of,  255 

flow  through,  254 

Junkers      two-cycle      fuel-injection 
engine,  440 


K 


Kerosene,  as  fuel  in  airplane  engines, 
436 


Lanchester  balancer,  58 

LeRhone  carburetor,  description  of, 

288 

29 


LeRhone  engine,  description  of,  185 
effect  of  compression  ratio  on 

horse  power,  393 
oil  pumps  of,  342 
performance  curves  of,  187 
Liberty  engine,  air  intake  for,  417 
battery     ignition    system    of, 

313 

breaker  mechanism  of,  318 
centrifugal  oil  cleaner  of,  333 
connecting  rods  of,  141 
cylinder  of,  124,  130 
description  of,  80 
dimensions  of,  (table),  72 
distributors  of,  318 
electric    starting    system    for, 

432 

generator,  315 
heat  balance,  39 
indicator  card,  40 
inertia  forces  and  bearing  loads, 

59 
influence    of    air    density    on 

performance  of,  388 
intake  manifolds  of,  419 
lubrication  system  of,  334 
method  of  starting,  425 
oil  pumps  of,  339 
performance  curves  of,  31,  83, 

239 

piston  of,  138 
starting  torque  of,  427 
torque  of,  43 

valve  action  of,  (curves),  168 
valve  of,  172 
water  pumps  of,  371 
weights  of  parts,  (table),  78 
Lift,  on  a  wing,  1 
Lubricating     oils,     properties     of, 

(table),  331 
reclaiming  of,  332 
specifications  for,  330 
tests  of,  330 
viscosity  of,  329 
Lubrication,  327-343 
methods  of,  333 
of  cylinders,  328 
of  radial  engines,  211 
oil  consumption  for,  338 


450 


INDEX 


M 


Magneto,  297-307 
armature  flux,  308 
typical  constants  for,  307 
for  unequal  firing  intervals,  300 
inductor  type,  300 
secondary  voltage  in,  309 
speed  of,  306 
starting,  425 
Manifolds,  exhaust,  421 
ejector  effect  in,  422 
intake,  417-421 

influence  of,  on  engine  perform- 
ance, 436 
preheating  of,  419 
pressure  drop  in,  28 
temperature  change  in,  418 
Master  carburetor,    description  of, 

283 
Materials,  for  special  engine  parts 

(table),  118 

properties  of,  (table),  114 
Maybach  carburetor,  description  of, 

286 

engine,  cylinder  of,  129 
dimensions  of,  124 
description  of,  113 
dimensions  of,  (table),  73 
.   fuel  pump  of,  293 
lubrication  system  of,  336 
performance  curves  of,  113 
pistons  of,  136 
starting  mechanism  of,  426 
valve  action  of,  (curves),  167 
valve  gear  of,  166 
water  pumps  of,  374 
weights  of  parts,  (table)  78 
Mechanical  efficiency,  23 
Mean  effective  pressure,  (table)  67 
brake,  25 
maximum     obtainable,     237, 

238 
Mercedes     engine,    air    pumps    of, 

292 

dimensions  of  (table),  73 
weights  of  parts  (table),  78 
Miller    carburetor,    description    of, 
282 


Mixture  strength   (see  also  Air-fuel 

Ratio) 

influence  of,  on  engine  perform- 
ance, 22,  32 

Moss  turbo-supercharger,  411 

Mufflers,,  exhaust  gas,  422 


X 


Napier  "Lion, "  connecting  rods  of, 

142 

description  of,  103 
lubrication  system  of,  335 
performance  curves  of,  106 


Odier  portable  starter,  433 
Offset  cylinders,  48 
Oil  pumps,  336,  338 

performance  curves  of,  340 

weights  of  (table),  78 
Oil  sumps,  150 
Oil  tanks,  343 
Oversized  engines,  390-394 
Orifice,    discharge    coefficients    for 

sharp-edged,  243 
Otto  cycle,  13 

efficiency  with  variable  specific 

heat,  19,  21 

Oxygen,  use  in  engine  at  altitudes, 
413 


Packard  engine,  cylinder  head  of, 

131 

description  of,  80 
dimensions  of  (table),  72 
performance  curves  of,  84 
weights  of  parts  of  (table),  78 

Parasite  resistance,  3 

Performance  curves,   Benz   engine, 

111 

Bugatti  engine,  98 
Clerget  engine,  190 
Cosmos  "Jupiter"  engine,  196 
Curtiss  engine,  90,  92 
Hall-Scott  engine,  98 


INDEX 


451 


Performance  curves,  Hispano-Suiza 

engine,  86 

LeRhone  engine,  187 
Liberty  engine,  83 
Maybach  engine,  113 
Napier  "Lion"  engine,  106 
Packard  engine,  84 
Salmson  engine,  199 
Siddeley  "Puma"  engine,  106 
Rolls-Royce  engine,  102 
Pipes,  for  fuel  system,  290 
Piston,      Pistons,      dimensions     of, 

(table),  72 
displacement,      per     horsepower, 

(table),  67 
divided-skirt,  135 
friction  of,  133 
material  of,  132 
pin,  138 

dimensions  of,  (table),  74 
loads  on,  (table),  76 
rings,  137 

dimension  of,  74 
slap  of,  134. 
slipper,  133 
speed, (table),  66 
weights  of,  (table),  76,  78,  166 
working  temperatures  of,  132 
Pitch,    of  propeller,    (see    Propeller 

Pitch  of) 

Power,  (see  Horsepower  and  Capacity) 
available,  for  flight,  8 
required,  for  flight,  1 
Priming,  of  engine,  423 
Pumps,  fuel,  289,  292 
Pressure,  Pressures,  in   ideal  (air) 

cycle,  14 
drop,  in  intake  manifold,  28 

past  valves,  152 
mean  effective,  25 
on  bearings,  144 
on  crankpin,  43 
on  piston,  40 
Propeller,  5 
coefficients,  7 
efficiency  of,  6 
geared,      (see      Geared     Propeller 

Drives) 
hubs,  147 


Propeller  hubs,  weights  of,  (table),  78 
pitch,  6 
pitch  ratio,  6 
slip,  6 

speed,  (table),  66 
thrust,  6 

thrust  horse  power  of,  6 
torque,  6 
torque  horse  power  of,  6 

R 

Radial  engines,  176,  193 

air-cooled,        dimensions       of, 

(table),  66 
ball  and  roller  crankpin  bearings 

in,  207 

crankshaft  of,  211 
details  of,  200 
firing  order  of,  178 
lubrication  of,  211 
number  of  cylinders  of,  178 
overall  dimensions  of,  (table),  79 
unbalanced  forces  in,  56 
valve  operation  in,  211 
water-cooled,      dimensions     of, 

(table),  68 
Radiators,  351-370 

constants  of,  (table),  362 
construction  of  complete,  368 
cores  of,  351 

dimensions  of,  359 
effect  of  position  on  performance 

of,  363 

figure  of  merit  of,  355,  (table),  360 
head     resistance     of,     354,     356, 

(table),  360 
heat  transfer  in,  352,  354,  (table), 

360 
horse   power   absorbed    by,    354, 

(table),  360 

limiting  temperatures  in,  363 
masking  of,  366,  368 
mass  flow  of  air  in,  (def.),  352 
obstructed,  352 

on  lighter-than-air  machines,  367 
performance  of,  358,  360 

at  altitudes,  365 
resistance  to  water  flow  in,  364 
selection  of,  356 


452 


INDEX 


Radiators,  size  of,  361 
water  flow  through,  363 
yawing  of,  368 
Rateau  supercharger,  410 
Reduction  gearing,   (see  Geared  Pro- 
peller Drives) 
Renault  engine,  connecting  rods  of, 

142 

cylinder  dimensions  of,  124 
propeller  gears  for,  378 
Resistance,  parasite,  3 
of  plane,  4 
of  wing,  2 
Revolutions,     of     typical    engines, 

(table),  66 
Ricardo  system,   of    supercharging, 

401 

Rocker  arms,  171 

Rolls-Royce  engine,  description  of,  98 
dimensions  of,  (table),  101 
performance  curves  of,  102 
propeller  gears  for,  380,  382 
Rotary  engines,  176,  179-193 
balance  of,  57 
details  of,  200 
dimensions  of,  (table),  66 
firing  order  of,  178 
number  of  cylinders  of,  178 
overall  dimensions  of  (table),  79 
torque  of,  50 

S 

Salmson  engine,  description  of,  198 

performance  curves  of,  199 
Saturation  temperature  of  air-fuel 

mixtures,  (table),  231 
Side  thrust,  48 
Siddeley  "Puma"  engine,  11 
connecting  rod  of,  140 
description  of,  105 
performance  curves  of,  106 
valves  of,  173 

Siemens-Halske    double-rotary    en- 
gine, 192 

Slip,  of  propeller,  6 
Spark  advance,  (see  Ignition) 
gaps,  310 
plugs,  319-324 

causes  of  failure  of,  320 


Spark      plugs,      construction     of, 

322 

cracking  of  insulators  of,  321 
dimensions  of,  76 
location  and  number  of,  324 
Sparking  voltages,  311 
Specific  heats,  at  constant  volume, 

(table),  20 
Springs,  safe  loads  and  deflections 

of,  170 
Starting,  424-433 

air  compressors  for,  430 
by  hand,  424 
Christensen  system,  431 
compressed  air  for,  429 
compression  release  for,  427 
electric  systems  for,  432 
magneto,  425 

mechanisms,  dog  clutch  for,  428 
integral,  429 
safety  devices  for,  429 
Motor-Compressor  Company  sys- 
tem, 430 

portable  crankers  for,  433 
torque,  426 

use  of  hydrogen  for,  425 
Steam-power  plants,  443 
Steels,  116-117      . 

for  exhaust  valves,  160 
Stewart- Warner  carburetor,  air  dis- 
charge coefficients  of,  252 
mixture  characteristics  of,  258 
performance  curves  of,  265 
Storage  cells,   (see  also  Battery),  312 
Strainers,  for  fuel  systems,  290 
Stroke-bore  ratios,  66,  72 
Stromberg    carburetor,    description 

of,  278 
mixture       characteristics      of, 

(curves),  258 

performance  curves  of,  266 
Sturtevant  engine,  cylinder  of,  128 

supercharger  for,  404 
Supercharged  engines,  394-413 
explosion  relief  valve  for,  408 
fuel  supply  system  for,  412 
power  developed  by,  394 
relief  valve  of,  413 
throttle  valve  for,  408 


INDEX 


453 


Superchargers,   (see  also  Air  Com- 
pression) t  397 

Brown-Boveri,  407 

coupling  for,  407 

exhaust  gas  turbine,  398,  408 

Moss  turbo-,  411 

Rateau,  410 

Schwade,  406 

Sturtevant,  404 
Supercharging,  368-415 

centrifugal  compressors  for,  403 

efficiency  of,  compressor,  400 
exhaust  gas  turbine,  400 

gearing  of  compressors  for,  403 

influence  on  plane  performance, 
415 

methods,  401 

multi-stage  compressors  for,  404 

reciprocating  compressors  for,  403 

Ricardo  system  of,  401 

Roots  positive  blower  for,  403 


Tangential  factors,  (table),  44 
Tanks,  fuel,  290 

oil,  343 

Tappet  clearance,  adjustment  of,  172 
Temperatures,  in  ideal    (air)  cycle, 

14 

of  jacket-water,  influence  on  en- 
gine capacity,  37J 
Test  results,  correction  to  standard 

conditions,  32 
Thermal  efficiency,  25 

effect  of  mixture  strength  on,  22 
maximum  observed,  21 
variation  with  air  temperature, 

35 

mixture  strength,  33 
throttling,  36 
Throttling,  influence  of,  on  thermal 

efficiency,  36 
Thrust,  of  propeller,  1,  6 
Timing  (see  Valve  Timing} 
Toluene,  effect  on  detonating  tem- 
perature, (table),  239 
value  (def.),  238 
Torque,  at  engine  crankshaft,  43 


Torque,  in  rotary  engines,  50 
of  engine,  26,  67 

at  starting,  426 
of  propeller,  6 

ratio  of  maximum  to  mean,  46 
variation  with  number  of  cylin- 
ders (table),  47 
Torsional  vibration,  of  crank-shaft, 

146 

Turbines,  exhaust  gas,  (see  Exhaust 
Gas  Turbines,  Super- 
charging) 

Turbo-superchargers,      (see     Super- 
chargers) 
Turbulence,  235 
Turning  moment,  (see  also  Torque), 

40 
Two-cycle  engines,  439 

U 

Unbalanced  forces,  magnitude  of,  56 
neutralizing  of,  58 
periodic,  effects  of,  57 


Valve,  Valves,  automatic  throttle,  391 
causes  of  failure  of,  160 
effect    of    lift    and    diameter   on 

engine  capacity,  158 
exhaust,  156 

dimensions  of,  66,  72 

temperature    of,    in    air-cooled 

cylinder,  349 
gears,  162 

weights  of,  (table),  78 
inlet,  dimensions  of,  66,  72 

number  per  cylinder,  66,  72 
lift  of,  151 

materials  of,  159,  161 
operation,  in  radial  engines,  211 
ports,  125 

pressure  drop  past,  152 
seats,  125 
springs,  168 

retainers  for,  170 

tension  of,  (table),  72 
stem  guides,  125 
temperatures  of,  159 
timing  of,  172,  173,  175 


454 


INDEX 


Valves,  weights  of,  (table),  76 
Variable-compression  engine,  438 

stroke  engine,  437 
Vee  engines,  angle  of,  64 
connecting  rods  of,  139 
dimensions,  (table),  70 
•     overall  dimensions  of,  (table),  79 

unbalanced  forces  in,  56 
Venturi  tube,  251 

discharge  coefficients  of,  252 
Vertical    engines,     dimensions    of, 

(table),  68 

overall  dimensions  of  (table),  79 
unbalanced  forces  in,  56 
Vibration,  (see  Torsional  Vibration) 
Viscosimeter,  329 
Viscosity,  measurement  of,  329 
of  fuels,  257 
of  oils,  329 

Volatility,  (see  also  Distillation),  220 
Volumetric  efficiency,  27,  28 

W 

W  engines,  arrangements  of,  65 
dimensions  of,  (table),  70 


Water  cooling,  350 
pumps,  370 

horse  power  required  for,  370 
performance  curves  of,  373,  374 
weights  of,  (table),  78 
vapor,  dissociation  of,  20 
Weights,  of  engines,  60,  67,  78 
of  engine  parts,  (table),  78 
of     reciprocating    parts,    (table), 

76 

of  rotating  parts,  (table),  76 
of  water  in  engine,  (table),  78 
Wings,  characteristics  of,  2 
Wiring  systems,  325 
Wright  engine,  description  of,  84 
Wrist  pin,  (see  Piston  pin) 


Zeitlen  engine,  description  of,  437 
Zenith     carburetor,     air     discharge 

coefficients  of,  252 
description  of,  272 
mixture  characteristics  of,  258 
performance  curves  of,  264 


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